Reliability of unstable periodic orbit based control strategies in biological systems

NASA Astrophysics Data System (ADS)

Mishra, Nagender; Hasse, Maria; Biswal, B.; Singh, Harinder P.

2015-04-01

Presence of recurrent and statistically significant unstable periodic orbits (UPOs) in time series obtained from biological systems is now routinely used as evidence for low dimensional chaos. Extracting accurate dynamical information from the detected UPO trajectories is vital for successful control strategies that either aim to stabilize the system near the fixed point or steer the system away from the periodic orbits. A hybrid UPO detection method from return maps that combines topological recurrence criterion, matrix fit algorithm, and stringent criterion for fixed point location gives accurate and statistically significant UPOs even in the presence of significant noise. Geometry of the return map, frequency of UPOs visiting the same trajectory, length of the data set, strength of the noise, and degree of nonstationarity affect the efficacy of the proposed method. Results suggest that establishing determinism from unambiguous UPO detection is often possible in short data sets with significant noise, but derived dynamical properties are rarely accurate and adequate for controlling the dynamics around these UPOs. A repeat chaos control experiment on epileptic hippocampal slices through more stringent control strategy and adaptive UPO tracking is reinterpreted in this context through simulation of similar control experiments on an analogous but stochastic computer model of epileptic brain slices. Reproduction of equivalent results suggests that far more stringent criteria are needed for linking apparent success of control in such experiments with possible determinism in the underlying dynamics.

Guidance and control strategies for aerospace vehicles

NASA Technical Reports Server (NTRS)

Naidu, Desineni S.; Hibey, Joseph L.

1988-01-01

The optimal control problem arising in coplanar, orbital transfer employing aeroassist technology is addressed. The maneuver involves the transfer from high Earth orbit to low Earth orbit. A performance index is chosen the minimize the fuel consumpltion for the transfer. Simulations are carried out for establishing a corridor of entry conditions which are suitable for flying the spacecraft through the atmosphere. A highlight of the paper is the application of an efficient multiple shooting method for taming the notorious nonlinear, two-point, boundary value problem resulting from optimization procedure.

An interplanetary targeting and orbit insertion maneuver design technique

NASA Technical Reports Server (NTRS)

Hintz, G. R.

1980-01-01

The paper describes a tradeoff in selecting a planetary encounter aimpoint and a spacecraft propulsive maneuver strategy in the Pioneer Venus Orbiter Mission. The method uses parametric data spanning a region of acceptable targeting aimpoints in the delivery space and the geometric considerations. Real-time maneuver adjustments accounted for known attitude control errors, orbit determination updates, and late changes in a targeting specification.

Autonomous Mars ascent and orbit rendezvous for earth return missions

NASA Technical Reports Server (NTRS)

Edwards, H. C.; Balmanno, W. F.; Cruz, Manuel I.; Ilgen, Marc R.

1991-01-01

The details of tha assessment of autonomous Mars ascent and orbit rendezvous for earth return missions are presented. Analyses addressing navigation system assessments, trajectory planning, targeting approaches, flight control guidance strategies, and performance sensitivities are included. Tradeoffs in the analysis and design process are discussed.

Optimal Control Strategies for Constrained Relative Orbits

DTIC Science & Technology

2007-09-01

the chief. The work assumes the Clohessy - Wiltshire closeness assump- tion between the deputy and chief is valid, however, elliptical chief orbits are...133 Appendix G. A Closed-Form Solution of the Linear Clohessy - Wiltshire Equa- tions...Counterspace . . . . . . . . . . . . . . . . . . . 1 CW Clohessy - Wiltshire . . . . . . . . . . . . . . . . . . . . . . 4 DARPA Defense Advanced Research

Modeling the Exo-Brake and the Development of Strategies for De-Orbit Drag Modulation

NASA Technical Reports Server (NTRS)

Murbach, M. S.; Papadopoulos, P.; Glass, C.; Dwyer-Cianciolo, A.; Powell, R. W.; Dutta, S.; Guarneros-Luna, A.; Tanner, F. A.; Dono, A.

2016-01-01

The Exo-Brake is a simple, non-propulsive means of de-orbiting small payloads from orbital platforms such as the International Space Station (ISS). Two de-orbiting experiments with fixed surface area Exo-Brakes have been successfully conducted in the last two years on the TechEdSat-3 and -4 nano-satellite missions. The development of the free molecular flow aerodynamic data-base is presented in terms of angle of attack, projected front surface area variation, and altitude. Altitudes are considered ranging from the 400km ISS jettison altitude to 90km. Trajectory tools are then used to predict de-orbit/entry corridors with the inclusion of the key atmospheric and geomagnetic uncertainties. Control system strategies are discussed which will be applied to the next two planned TechEdSat-5 and -6 nano-satellite missions - thus increasing the targeting accuracy at the Von Karman altitude through the proposed drag modulation technique.

Reliability of unstable periodic orbit based control strategies in biological systems

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

Mishra, Nagender; Singh, Harinder P.; Hasse, Maria

2015-04-15

Presence of recurrent and statistically significant unstable periodic orbits (UPOs) in time series obtained from biological systems is now routinely used as evidence for low dimensional chaos. Extracting accurate dynamical information from the detected UPO trajectories is vital for successful control strategies that either aim to stabilize the system near the fixed point or steer the system away from the periodic orbits. A hybrid UPO detection method from return maps that combines topological recurrence criterion, matrix fit algorithm, and stringent criterion for fixed point location gives accurate and statistically significant UPOs even in the presence of significant noise. Geometry ofmore » the return map, frequency of UPOs visiting the same trajectory, length of the data set, strength of the noise, and degree of nonstationarity affect the efficacy of the proposed method. Results suggest that establishing determinism from unambiguous UPO detection is often possible in short data sets with significant noise, but derived dynamical properties are rarely accurate and adequate for controlling the dynamics around these UPOs. A repeat chaos control experiment on epileptic hippocampal slices through more stringent control strategy and adaptive UPO tracking is reinterpreted in this context through simulation of similar control experiments on an analogous but stochastic computer model of epileptic brain slices. Reproduction of equivalent results suggests that far more stringent criteria are needed for linking apparent success of control in such experiments with possible determinism in the underlying dynamics.« less

Integrated Attitude Control Strategy for the Asteroid Redirect Mission

NASA Technical Reports Server (NTRS)

Lopez, Pedro, Jr.; Price, Hoppy; San Martin, Miguel

2014-01-01

A deep-space mission has been proposed to redirect an asteroid to a distant retrograde orbit around the moon using a robotic vehicle, the Asteroid Redirect Vehicle (ARV). In this orbit, astronauts will rendezvous with the ARV using the Orion spacecraft. The integrated attitude control concept that Orion will use for approach and docking and for mated operations will be described. Details of the ARV's attitude control system and its associated constraints for redirecting the asteroid to the distant retrograde orbit around the moon will be provided. Once Orion is docked to the ARV, an overall description of the mated stack attitude during all phases of the mission will be presented using a coordinate system that was developed for this mission. Next, the thermal and power constraints of both the ARV and Orion will be discussed as well as how they are used to define the optimal integrated stack attitude. Lastly, the lighting and communications constraints necessary for the crew's extravehicular activity planned to retrieve samples from the asteroid will be examined. Similarly, the joint attitude control strategy that employs both the Orion and the ARV attitude control assets prior, during, and after each extravehicular activity will also be thoroughly discussed.

Guidance, Navigation, and Control Techniques and Technologies for Active Satellite Removal

NASA Astrophysics Data System (ADS)

Ortega Hernando, Guillermo; Erb, Sven; Cropp, Alexander; Voirin, Thomas; Dubois-Matra, Olivier; Rinalducci, Antonio; Visentin, Gianfranco; Innocenti, Luisa; Raposo, Ana

2013-09-01

This paper shows an internal feasibility analysis to de- orbit a non-functional satellite of big dimensions by the Technical Directorate of the European Space Agency ESA. The paper focuses specifically on the design of the techniques and technologies for the Guidance, Navigation, and Control (GNC) system of the spacecraft mission that will capture the satellite and ultimately will de-orbit it on a controlled re-entry.The paper explains the guidance strategies to launch, rendezvous, close-approach, and capture the target satellite. The guidance strategy uses chaser manoeuvres, hold points, and collision avoidance trajectories to ensure a safe capture. It also details the guidance profile to de-orbit it in a controlled re-entry.The paper continues with an analysis of the required sensing suite and the navigation algorithms to allow the homing, fly-around, and capture of the target satellite. The emphasis is placed around the design of a system to allow the rendezvous with an un-cooperative target, including the autonomous acquisition of both the orbital elements and the attitude of the target satellite.Analysing the capture phase, the paper provides a trade- off between two selected capture systems: the net and the tentacles. Both are studied from the point of view of the GNC system.The paper analyses as well the advanced algorithms proposed to control the final compound after the capture that will allow the controlled de-orbiting of the assembly in a safe place in the Earth.The paper ends proposing the continuation of this work with the extension to the analysis of the destruction process of the compound in consecutive segments starting from the entry gate to the rupture and break up.

Optimal aeroassisted orbital transfer with plane change using collocation and nonlinear programming

NASA Technical Reports Server (NTRS)

Shi, Yun. Y.; Nelson, R. L.; Young, D. H.

1990-01-01

The fuel optimal control problem arising in the non-planar orbital transfer employing aeroassisted technology is addressed. The mission involves the transfer from high energy orbit (HEO) to low energy orbit (LEO) with orbital plane change. The basic strategy here is to employ a combination of propulsive maneuvers in space and aerodynamic maneuvers in the atmosphere. The basic sequence of events for the aeroassisted HEO to LEO transfer consists of three phases. In the first phase, the orbital transfer begins with a deorbit impulse at HEO which injects the vehicle into an elliptic transfer orbit with perigee inside the atmosphere. In the second phase, the vehicle is optimally controlled by lift and bank angle modulations to perform the desired orbital plane change and to satisfy heating constraints. Because of the energy loss during the turn, an impulse is required to initiate the third phase to boost the vehicle back to the desired LEO orbital altitude. The third impulse is then used to circularize the orbit at LEO. The problem is solved by a direct optimization technique which uses piecewise polynomial representation for the state and control variables and collocation to satisfy the differential equations. This technique converts the optimal control problem into a nonlinear programming problem which is solved numerically. Solutions were obtained for cases with and without heat constraints and for cases of different orbital inclination changes. The method appears to be more powerful and robust than other optimization methods. In addition, the method can handle complex dynamical constraints.

1999 Flight Mechanics Symposium

NASA Technical Reports Server (NTRS)

Lynch, John P. (Editor)

1999-01-01

This conference publication includes papers and abstracts presented at the Flight Mechanics Symposium held on May 18-20, 1999. Sponsored by the Guidance, Navigation and Control Center of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Model predictive control for spacecraft rendezvous in elliptical orbit

NASA Astrophysics Data System (ADS)

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.

Codimension-Two Bifurcation, Chaos and Control in a Discrete-Time Information Diffusion Model

NASA Astrophysics Data System (ADS)

Ren, Jingli; Yu, Liping

2016-12-01

In this paper, we present a discrete model to illustrate how two pieces of information interact with online social networks and investigate the dynamics of discrete-time information diffusion model in three types: reverse type, intervention type and mutualistic type. It is found that the model has orbits with period 2, 4, 6, 8, 12, 16, 20, 30, quasiperiodic orbit, and undergoes heteroclinic bifurcation near 1:2 point, a homoclinic structure near 1:3 resonance point and an invariant cycle bifurcated by period 4 orbit near 1:4 resonance point. Moreover, in order to regulate information diffusion process and information security, we give two control strategies, the hybrid control method and the feedback controller of polynomial functions, to control chaos, flip bifurcation, 1:2, 1:3 and 1:4 resonances, respectively, in the two-dimensional discrete system.

Orbital Debris: A Policy Perspective

NASA Technical Reports Server (NTRS)

Johnson, Nicholas L.

2007-01-01

A viewgraph presentation describing orbital debris from a policy perspective is shown. The contents include: 1) Voyage through near-Earth Space-animation; 2) What is Orbital Debris?; 3) Orbital Debris Detectors and Damage Potential; 4) Hubble Space Telescope; 5) Mir Space Station Solar Array; 6) International Space Station; 7) Space Shuttle; 8) Satellite Explosions; 9) Satellite Collisions; 10) NASA Orbital Debris Mitigation Guidelines; 11) International Space Station Jettison Policy; 12) Controlled/Uncontrolled Satellite Reentries; 13) Return of Space Objects; 14) Orbital Debris and U.S. National Space Policy; 15) U.S Government Policy Strategy; 16) Bankruptcy of the Iridium Satellite System; 17) Inter-Agency Space Debris Coordination Committee (IADC); 18) Orbital Debris at the United Nations; 19) Chinese Anti-satellite System; 20) Future Evolution of Satellite Population; and 21) Challenge of Orbital Debris

Coupled Attitude and Orbit Dynamics and Control in Formation Flying Systems

NASA Technical Reports Server (NTRS)

Xu, Yun-Jun; Fitz-Coy, Norman; Mason, Paul

2003-01-01

Formation flying systems can range from global constellations offering extended service coverage to clusters of highly coordinated vehicles that perform distributed sensing. Recently, the use of groups of micro-satellites in the areas of near Earth explorations, deep space explorations, and military applications has received considerable attention by researchers and practitioners. To date, most proposed control strategies are based on linear models (e.g., Hill-Clohessy-Wiltshire equations) or nonlinear models that are restricted to circular reference orbits. Also, all models in the literature are uncoupled between relative position and relative attitude. In this paper, a generalized dynamic model is proposed. The reference orbit is not restricted to the circular case. In this formulation, the leader or follower satellite can be in either a circular or an elliptic orbit. In addition to maintaining a specified relative position, the satellites are also required to maintain specified relative attitudes. Thus the model presented couples vehicle attitude and orbit requirements. Orbit perturbations are also included. In particular, the J(sub 2) effects are accounted in the model. Finally, a sliding mode controller is developed and used to control the relative attitude of the formation and the simulation results are presented.

Flight Mechanics Symposium 1997

NASA Technical Reports Server (NTRS)

Walls, Donna M. (Editor)

1997-01-01

This conference publication includes papers and abstracts presented at the Flight Mechanics Symposium. This symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Touring the Saturnian system

NASA Astrophysics Data System (ADS)

Wolf, Aron A.

1996-10-01

The Cassini mission to Saturn employs a Saturn orbiter and a Titan probe to conduct an intensive investigation of the Saturnian system. The Cassini orbiter flies a series of obits, incorporating flybys of the Saturnian satellites, called the 'satellite tour'. During the tour, the gravitational fields of the satellites are used to modify and control the orbit, targeting from one satellite flyby to the next. The tour trajectory must also be designed to maximize opportunities for science observations, subject to mission-imposed constraints. Tour design studies have been conducted for Cassini to identify trades and strategies for achieving these sometimes conflicting goals. Concepts, strategies, and techniques previously developed for the Galileo mission to Jupiter have been modified, and new ones have been developed, to meet the requirements of the Cassini mission.

Air-breathing hypersonic vehicle guidance and control studies; An integrated trajectory/control analysis methodology: Phase 1

NASA Technical Reports Server (NTRS)

Hattis, Philip D.; Malchow, Harvey L.

1991-01-01

A tool which generates optimal trajectory/control histories in an integrated manner is generically adapted to the treatment of single-stage-to-orbit air-breathing hypersonic vehicles. The methodology is implemented as a two point boundary value problem solution technique. Its use permits an assessment of an entire near-minimum-fuel trajectory and desired control strategy from takeoff to orbit while satisfying physically derived inequality constraints and while achieving efficient propulsive mode phasing. A simpler analysis strategy that partitions the trajectory into several boundary condition matched segments is also included to construct preliminary trajectory and control history representations with less computational burden than is required for the overall flight profile assessment. A demonstration was accomplished using a tabulated example (winged-cone accelerator) vehicle model that is combined with a newly developed multidimensional cubic spline data smoothing routine. A constrained near-fuel-optimal trajectory, imposing a dynamic pressure limit of 1000 psf, was developed from horizontal takeoff to 20,000 ft/sec relative air speed while aiming for a polar orbit. Previously unspecified propulsive discontinuities were located. Flight regimes demanding rapid attitude changes were identified, dictating control effector and closed-loop controller authority was ascertained after evaluating effector use for vehicle trim. Also, inadequacies in vehicle model representations and specific subsystem models with insufficient fidelity were determined based on unusual control characteristics and/or excessive sensitivity to uncertainty.

Analysis of Envisat Orbit Maintenance Strategies to Improve/Increase Envisat ASAR Interferometry Opportunities

NASA Technical Reports Server (NTRS)

Kuijper, D.; Matatoros, Garcia

2007-01-01

The biggest and most advanced Earth Observation Satellite in-orbit, developed by the European Space Agency (ESA) and its member states, is Envisat. It was launched on March 1, 2002 by an Ariane V from French Guyana and holds a total of 10 multi-disciplinary Earth observation instruments, among which an Advanced Synthetic Aperture Radar (ASAR). The ASAR user community requested the Flight Dynamics division of the European Space Operations Centre (ESOC) to investigate how the orbit control maintenance strategy for Envisat could be changed to optimize ASAR interferometry opportunities overall and in addition support the International Polar Year 2007/2008 initiative. The Polar Regions play a pivotal role in understanding our planet and our impact on it as they are recognized as sensitive barometers of environmental change. One of the main themes of the International Polar Year 2007/2008 is therefore the study of Earth s changing ice and snow, and its impact on our planet and our lives. Naturally, ESA would like to support this very important initiative. This paper presents the investigations that have been conducted to support these requests in the best possible way. It discusses the orbit maintenance strategy that has been in place since its launch, ensuring the actual orbit to be within 1 km of a so-called reference orbit, and presents the new orbit maintenance strategy that is aimed at improving/increasing the opportunities for Envisat ASAR interferometry, while preserving the fuel on board the spacecraft. The hydrazine on-board Envisat happens to be a precious resource as only approximately 300 kg of it was available at launch, like ERS-2. The difference being however that the mass of Envisat is approximately 3.2 times that of ERS-2.

State dependent model predictive control for orbital rendezvous using pulse-width pulse-frequency modulated thrusters

NASA Astrophysics Data System (ADS)

Li, Peng; Zhu, Zheng H.; Meguid, S. A.

2016-07-01

This paper studies the pulse-width pulse-frequency modulation based trajectory planning for orbital rendezvous and proximity maneuvering near a non-cooperative spacecraft in an elliptical orbit. The problem is formulated by converting the continuous control input, output from the state dependent model predictive control, into a sequence of pulses of constant magnitude by controlling firing frequency and duration of constant-magnitude thrusters. The state dependent model predictive control is derived by minimizing the control error of states and control roughness of control input for a safe, smooth and fuel efficient approaching trajectory. The resulting nonlinear programming problem is converted into a series of quadratic programming problem and solved by numerical iteration using the receding horizon strategy. The numerical results show that the proposed state dependent model predictive control with the pulse-width pulse-frequency modulation is able to effectively generate optimized trajectories using equivalent control pulses for the proximity maneuvering with less energy consumption.

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

NASA Technical Reports Server (NTRS)

Conn, Tracie R.; 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 fuel 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 delta-v required for long-term station keeping and minimummaximummean spacecraft separation distances are also presented.

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

NASA Astrophysics Data System (ADS)

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.

Estimating maneuvers for precise relative orbit determination using GPS

NASA Astrophysics Data System (ADS)

Allende-Alba, Gerardo; Montenbruck, Oliver; Ardaens, Jean-Sébastien; Wermuth, Martin; Hugentobler, Urs

2017-01-01

Precise relative orbit determination is an essential element for the generation of science products from distributed instrumentation of formation flying satellites in low Earth orbit. According to the mission profile, the required formation is typically maintained and/or controlled by executing maneuvers. In order to generate consistent and precise orbit products, a strategy for maneuver handling is mandatory in order to avoid discontinuities or precision degradation before, after and during maneuver execution. Precise orbit determination offers the possibility of maneuver estimation in an adjustment of single-satellite trajectories using GPS measurements. However, a consistent formulation of a precise relative orbit determination scheme requires the implementation of a maneuver estimation strategy which can be used, in addition, to improve the precision of maneuver estimates by drawing upon the use of differential GPS measurements. The present study introduces a method for precise relative orbit determination based on a reduced-dynamic batch processing of differential GPS pseudorange and carrier phase measurements, which includes maneuver estimation as part of the relative orbit adjustment. The proposed method has been validated using flight data from space missions with different rates of maneuvering activity, including the GRACE, TanDEM-X and PRISMA missions. The results show the feasibility of obtaining precise relative orbits without degradation in the vicinity of maneuvers as well as improved maneuver estimates that can be used for better maneuver planning in flight dynamics operations.

Simulator evaluation of display concepts for pilot monitoring and control of space shuttle approach and landing. Phase 2: Manual flight control

NASA Technical Reports Server (NTRS)

Gartner, W. B.; Baldwin, K. M.

1973-01-01

A study of the display requirements for final approach management of the space shuttle orbiter vehicle is presented. An experimental display concept, providing a more direct, pictorial representation of the vehicle's movement relative to the selected approach path and aiming points, was developed and assessed as an aid to manual flight path control. Both head-up, windshield projections and head-down, panel mounted presentations of the experimental display were evaluated in a series of simulated orbiter approach sequence. Data obtained indicate that the experimental display would enable orbiter pilots to exercise greater flexibility in implementing alternative final approach control strategies. Touchdown position and airspeed dispersion criteria were satisfied on 91 percent of the approach sequences, representing various profile and wind effect conditions. Flight path control and airspeed management satisfied operationally-relevant criteria for the two-segment, power-off orbiter approach and were consistently more accurate and less variable when the full set of experimental display elements was available to the pilot. Approach control tended to be more precise when the head-up display was used; however, the data also indicate that the head-down display would provide adequate support for the manual control task.

Orbit Determination Issues for Libration Point Orbits

NASA Technical Reports Server (NTRS)

Beckman, Mark; Bauer, Frank (Technical Monitor)

2002-01-01

Libration point mission designers require knowledge of orbital accuracy for a variety of analyses including station keeping control strategies, transfer trajectory design, and formation and constellation control. Past publications have detailed orbit determination (OD) results from individual libration point missions. This paper collects both published and unpublished results from four previous libration point missions (ISEE (International Sun-Earth Explorer) -3, SOHO (Solar and Heliospheric Observatory), ACE (Advanced Composition Explorer) and MAP (Microwave Anisotropy Probe)) supported by Goddard Space Flight Center's Guidance, Navigation & Control Center. The results of those missions are presented along with OD issues specific to each mission. All past missions have been limited to ground based tracking through NASA ground sites using standard range and Doppler measurement types. Advanced technology is enabling other OD options including onboard navigation using seaboard attitude sensors and the use of the Very Long Baseline Interferometry (VLBI) measurement Delta Differenced One-Way Range (DDOR). Both options potentially enable missions to reduce coherent dedicated tracking passes while maintaining orbital accuracy. With the increased projected loading of the DSN (Deep Space Network), missions must find alternatives to the standard OD scenario.

An introduction to shuttle/LDEF retrieval operations: The R-bar approach option. [orbital mechanics and braking schedule

NASA Technical Reports Server (NTRS)

Hall, W. M.

1978-01-01

Simulated orbiter direct approaches during long duration exposure facility (LDEF) retrieval operations reveal that the resultant orbiter jet plume fields can significantly disturb LDEF. An alternate approach technique which utilizes orbital mechanics forces in lieu of jets to brake the final orbiter/LDEF relative motion during the final approach, is described. Topics discussed include: rendezvous operations from the terminal phase initiation burn through braking at some standoff distance from LDEF, pilot and copilot activities, the cockpit instrumentation employed, and a convenient coordinate frame for studying the relative motion between two orbiting bodies. The basic equations of motion for operating on the LDEF radius vector are introduced. Practical considerations of implementing an R-bar approach, namely, orbiter/LDEF relative state uncertainties and orbiter control system limitations are explored. A possible R-bar approach strategy is developed and demonstrated.

Primeval substance delivery from Phobos to the Earth—the Phobos-Soil project: Ballistics, navigation, and flight control

NASA Astrophysics Data System (ADS)

Akim, E. L.; Zaslavsky, G. S.; Morskoy, I. M.; Ruzsky, E. G.; Stepaniants, V. A.; Tuchin, A. G.

2010-02-01

This paper is concerned with the problems of ballistics, navigation, and flight control of the space craft (SC) in the Phobos-Grunt mission. We consider an insertion into the Earth-Mars transfer trajectory, the Earth-Mars transfer, the strategy of corrections, and the accuracy of the insertion of the SC into Martian orbit. During the orbital maneuvering stage in the sphere of influence of Mars, we set up a scheme that allows for the insertion of the SC, with the prescribed accuracy, into a point 80-km above the Phobos surface over the theoretical landing area. We specify the sequence for a controlled landing and provide methods for solving the problems of navigation and control during a self-c ontained landing. We also consider the liftoff from Phobos, insertion into the parking orbit, and the Mars-Earth transfer.

Cassini Orbit Trim Maneuvers at Saturn - Overview of Attitude Control Flight Operations

NASA Technical Reports Server (NTRS)

Burk, Thomas A.

2011-01-01

The Cassini spacecraft has been in orbit around Saturn since July 1, 2004. To remain on the planned trajectory which maximizes science data return, Cassini must perform orbit trim maneuvers using either its main engine or its reaction control system thrusters. Over 200 maneuvers have been executed on the spacecraft since arrival at Saturn. To improve performance and maintain spacecraft health, changes have been made in maneuver design command placement, in accelerometer scale factor, and in the pre-aim vector used to align the engine gimbal actuator prior to main engine burn ignition. These and other changes have improved maneuver performance execution errors significantly since 2004. A strategy has been developed to decide whether a main engine maneuver should be performed, or whether the maneuver can be executed using the reaction control system.

Four-Formation In-Track Configuration Maintenance Strategy

NASA Technical Reports Server (NTRS)

Lamy, Alain; Costes, Thierry

2007-01-01

The aim of this paper is to present the analysis conducted by CNES for the maintenance of a formation made of several LEO satellites (typically 4) in several planes (typically 2), 100 km or so apart from each other. The along-track separations between the satellites have to be controlled to within 15 km thanks to orbit correction maneuvers supposed to be performed every 2 weeks. The main difficulty is related to solar activity which is expected to be close to its maximum for the entire mission s lifespan. As a matter of fact, a high solar activity makes orbit prediction harder, and makes it impossible to keep the altitude of the formation constant. Thus, a specific relative maintenance strategy had to be devised in order to meet the mission's requirements. The first part provides a few elements on the mission analysis process that has taken place. The method used for the evaluation of the maneuver frequency is detailed, based on the evaluation of the effects of atmospheric drag on the orbit. The second part is dedicated to the maintenance strategy that has been designed, and particularly to the computation of the reference orbits and of the velocity increments that enable the in-track inter-satellite distances to be maintained within the desired bounds. Finally a few simulation results are presented; they enable the performance of the maintenance strategy to be checked in a more realistic context.

On the magnetic attitude control for spacecraft via the ɛ-strategies method

NASA Astrophysics Data System (ADS)

Smirnov, Georgi V.; Ovchinnikov, Mikhail; Miranda, Francisco

2008-09-01

We develop a new approach to stabilization problems based on a combination of the Lyapunov functions method with local controllability properties. The stabilizability is understood in the sense of ɛ-strategies introduced by Pontryagin in the frame of differential games theory. To illustrate the possibilities of our approach we consider a satellite with two magnetic coils directed along its principal inertia axes. Its circular orbit is neither polar nor equatorial. We show that there exists an ɛ-strategy stabilizing an Earth pointing satellite, whenever the deviations from the equilibrium position are small enough.

Orbital management and design considerations for NiCd satellite power systems

NASA Technical Reports Server (NTRS)

Tausch, Benjamin J., II

1994-01-01

Several recently manufactured 50 and 60 ampere hour aerospace NiCd battery cell lots, produced by Gates Aerospace Batteries, are prone to premature on orbit performance degradation. The failure mechanism is cadmium migration, and the consequent development of soft shorts. A Martin Marietta Astronautics satellite program instituted an orbital management strategy for a set of these batteries that reduced the rate of degradation and brought the system to stable operation. This strategy involves: (1) minimizing the accumulated battery overcharge; (2) regular discharge exercises; and (3) periodic battery reconditioning. Because of changes in the NiCd cell manufacturing process, the actual performance of subsequent lots of NiCd cells is open to question. Future NiCd based power system designs should therefore allow for fine control of charge parameters, and an on orbit battery reconditioning capability. To minimize risk, it is much better to perform a full life test to qualify the cells before launch, rather than in parallel with orbital operations. If there are any changes in the manufacturing process of cells, it is extremely important to maintain very strong cognizance of secondary subcontractors, recognizing that the cell and battery manufacturing discipline is easily atrophied.

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

NASA Technical Reports Server (NTRS)

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.

Orbital atherectomy for treating de novo, severely calcified coronary lesions: 3-year results of the pivotal ORBIT II trial.

PubMed

Lee, Michael; Généreux, Philippe; Shlofmitz, Richard; Phillipson, Daniel; Anose, Bynthia M; Martinsen, Brad J; Himmelstein, Stevan I; Chambers, Jeff W

2017-06-01

The presence of heavy coronary artery calcification increases the complexity of percutaneous coronary intervention (PCI) and increases the incidence of major adverse cardiac events (MACE): death, myocardial infarction (MI), target vessel revascularization (TVR), and stent thrombosis. The ORBIT II (Evaluate the Safety and Efficacy of OAS in Treating Severely Calcified Coronary Lesions) trial reported low rates of procedural, 30-day, 1-year, and 2-year ischemic complications after treatment of de novo, severely calcified lesions with the Diamondback 360° Coronary Orbital Atherectomy System (OAS) (Cardiovascular Systems, Inc.). ORBIT II was a single-arm trial that enrolled 443 patients at 49U.S. sites; in this study, de novo, severely calcified coronary lesions were treated with OAS prior to stenting. The primary safety endpoint was 30-day MACE: the composite of cardiac death, MI, and TVR (inclusive of target lesion revascularization (TLR)). The primary efficacy endpoint was procedural success: stent delivery with a residual stenosis of <50% without the occurrence of in-hospital MACE.The present analysis reports the final, 3-year follow-up results from ORBIT II. The majority of subjects (88.2%) underwent PCI with drug-eluting stents after orbital atherectomy. There were 360 (81.3%) subjects who completed the protocol-mandated 3-year visit.The overall cumulative rate of 3-year MACE was 23.5%, including cardiac death (6.7%), MI (11.2%), and TVR (10.2%). The 3-year target lesion revascularization rate was 7.8%. In the final 3-year analysis of the ORBIT II trial, orbital atherectomy of severely calcified coronary lesions followed by stenting resulted in a low rate of adverse ischemic events compared with historical controls.Orbital atherectomy represents a safe and effective revascularization strategy for patients with severely calcified coronary lesions. The ORBIT II trial enrolled 443 subjects to study orbital atherectomy followed by stenting for de novo severely calcified coronary lesions. The overall cumulative 3-year MACE rate was 23.5%, including cardiac death (6.7%), MI (11.2%), and TVR (10.2%); the 3-year target lesion revascularization rate was 7.8%. Orbital atherectomy of heavily calcified coronary lesions followed by stenting results in a low rate of adverse ischemic events compared with historical controls; it represents a reasonable revascularization strategy for patients with severely calcified coronary lesions. Copyright © 2017 Elsevier Inc. All rights reserved.

Proceedings from the 2nd International Symposium on Formation Flying Missions and Technologies

NASA Technical Reports Server (NTRS)

2004-01-01

Topics discussed include: The Stellar Imager (SI) "Vision Mission"; First Formation Flying Demonstration Mission Including on Flight Nulling; Formation Flying X-ray Telescope in L2 Orbit; SPECS: The Kilometer-baseline Far-IR Interferometer in NASA's Space Science Roadmap Presentation; A Tight Formation for Along-track SAR Interferometry; Realization of the Solar Power Satellite using the Formation Flying Solar Reflector; SIMBOL-X : Formation Flying for High-Energy Astrophysics; High Precision Optical Metrology for DARWIN; Close Formation Flight of Micro-Satellites for SAR Interferometry; Station-Keeping Requirements for Astronomical Imaging with Constellations of Free-Flying Collectors; Closed-Loop Control of Formation Flying Satellites; Formation Control for the MAXIM Mission; Precision Formation Keeping at L2 Using the Autonomous Formation Flying Sensor; Robust Control of Multiple Spacecraft Formation Flying; Virtual Rigid Body (VRB) Satellite Formation Control: Stable Mode-Switching and Cross-Coupling; Electromagnetic Formation Flight (EMFF) System Design, Mission Capabilities, and Testbed Development; Navigation Algorithms for Formation Flying Missions; Use of Formation Flying Small Satellites Incorporating OISL's in a Tandem Cluster Mission; Semimajor Axis Estimation Strategies; Relative Attitude Determination of Earth Orbiting Formations Using GPS Receivers; Analysis of Formation Flying in Eccentric Orbits Using Linearized Equations of Relative Motion; Conservative Analytical Collision Probabilities for Orbital Formation Flying; Equations of Motion and Stability of Two Spacecraft in Formation at the Earth/Moon Triangular Libration Points; Formations Near the Libration Points: Design Strategies Using Natural and Non-Natural Ares; An Overview of the Formation and Attitude Control System for the Terrestrial Planet Finder Formation Flying Interferometer; GVE-Based Dynamics and Control for Formation Flying Spacecraft; GNC System Design for a New Concept of X-Ray Distributed Telescope; GNC System for the Deployment and Fine Control of the DARWIN Free-Flying Interferometer; Formation Algorithm and Simulation Testbed; and PLATFORM: A Formation Flying, RvD and Robotic Validation Test-bench.

A Minimum Delta V Orbit Maintenance Strategy for Low-Altitude Missions Using Burn Parameter Optimization

NASA Technical Reports Server (NTRS)

Brown, Aaron J.

2011-01-01

Orbit maintenance is the series of burns performed during a mission to ensure the orbit satisfies mission constraints. Low-altitude missions often require non-trivial orbit maintenance Delta V due to sizable orbital perturbations and minimum altitude thresholds. A strategy is presented for minimizing this Delta V using impulsive burn parameter optimization. An initial estimate for the burn parameters is generated by considering a feasible solution to the orbit maintenance problem. An low-lunar orbit example demonstrates the Delta V savings from the feasible solution to the optimal solution. The strategy s extensibility to more complex missions is discussed, as well as the limitations of its use.

Disposal strategy for the geosynchronous orbits of the Beidou Navigation Satellite System

NASA Astrophysics Data System (ADS)

Tang, Jingshi; Liu, Lin

Beidou Navigation Satellite System (BDS) is China's navigation satelite system. It is now operational for navigation service in China and Asia-Pacific region and is due to be fully operational as a global navigation system by 2020. Unlike other navigation satellite systems, BDS consists of both 12-hour medium Earth orbit and 24-hour geosynchronous orbit. To sustain a safe environment for the navigation satellites, the end-of-life satellites must be disposed appropriately so they do not pose potential dangers to the operational satellites. There are currently two strategies for the disposal orbit. One is to put the disposed satellite in a graveyard orbit that has a safe distance from the operational satellites. It is often applied in geosynchronous orbits and such graveyard orbit can always maintain a safe distance even for a few centuries. This strategy is also currently adopted by GPS, yet recent researches show a re-entry orbit can sometimes be a better alternative. The interaction of Earth oblateness and lunisolar gravitation can lead to a rapid increase in the orbit eccentricity such that by proper design the disposed GPS satellite can be cleared out by re-entry into the atmosphere. In this work we focus on the disposal strategy for BDS geosynchronous orbit, which consists of the equatorial stationary orbit (GEO) and the inclined orbit (IGSO). We show that these two orbits are essentially in two different dynamical environments and evolve quite distinctly over a long period of time. Taking advantage of the dynamic nature, we apply the graveyard orbit and the re-entry orbit to GEO and IGSO respectively and propose appropriate disposal strategies accordingly.

Orbital prefrontal cortex is required for object-in-place scene memory but not performance of a strategy implementation task.

PubMed

Baxter, Mark G; Gaffan, David; Kyriazis, Diana A; Mitchell, Anna S

2007-10-17

The orbital prefrontal cortex is thought to be involved in behavioral flexibility in primates, and human neuroimaging studies have identified orbital prefrontal activation during episodic memory encoding. The goal of the present study was to ascertain whether deficits in strategy implementation and episodic memory that occur after ablation of the entire prefrontal cortex can be ascribed to damage to the orbital prefrontal cortex. Rhesus monkeys were preoperatively trained on two behavioral tasks, the performance of both of which is severely impaired by the disconnection of frontal cortex from inferotemporal cortex. In the strategy implementation task, monkeys were required to learn about two categories of objects, each associated with a different strategy that had to be performed to obtain food reward. The different strategies had to be applied flexibly to optimize the rate of reward delivery. In the scene memory task, monkeys learned 20 new object-in-place discrimination problems in each session. Monkeys were tested on both tasks before and after bilateral ablation of orbital prefrontal cortex. These lesions impaired new scene learning but had no effect on strategy implementation. This finding supports a role for the orbital prefrontal cortex in memory but places limits on the involvement of orbital prefrontal cortex in the representation and implementation of behavioral goals and strategies.

Strategies for high-precision Global Positioning System orbit determination

NASA Technical Reports Server (NTRS)

Lichten, Stephen M.; Border, James S.

1987-01-01

Various strategies for the high-precision orbit determination of the GPS satellites are explored using data from the 1985 GPS field test. Several refinements to the orbit determination strategies were found to be crucial for achieving high levels of repeatability and accuracy. These include the fine tuning of the GPS solar radiation coefficients and the ground station zenith tropospheric delays. Multiday arcs of 3-6 days provided better orbits and baselines than the 8-hr arcs from single-day passes. Highest-quality orbits and baselines were obtained with combined carrier phase and pseudorange solutions.

Flight Mechanics/Estimation Theory Symposium, 1992

NASA Technical Reports Server (NTRS)

Stengle, Thomas H. (Editor)

1993-01-01

This conference publication includes 40 papers and abstracts presented at the Flight Mechanics/Estimation Theory Symposium on May 5-7, 1992. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Flight Mechanics/Estimation Theory Symposium 1996

NASA Technical Reports Server (NTRS)

Greatorex, Scott (Editor)

1996-01-01

This conference publication includes 34 papers and abstracts presented at the Flight Mechanics/ Estimation Theory Symposium on May 14-16, 1996. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Flight Mechanics/Estimation Theory Symposium, 1994

NASA Technical Reports Server (NTRS)

Hartman, Kathy R. (Editor)

1994-01-01

This conference publication includes 41 papers and abstracts presented at the Flight Mechanics/Estimation Theory Symposium on May 17-19, 1994. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Flight Mechanics/Estimation Theory Symposium, 1990

NASA Technical Reports Server (NTRS)

Stengle, Thomas (Editor)

1990-01-01

This conference publication includes 32 papers and abstracts presented at the Flight Mechanics/Estimation Theory Symposium on May 22-25, 1990. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium features technical papers on a wide range of issues related to orbit-attitude prediction, determination and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Flight Mechanics/Estimation Theory Symposium 1995

NASA Technical Reports Server (NTRS)

Hartman, Kathy R. (Editor)

1995-01-01

This conference publication includes 41 papers and abstracts presented at the Flight Mechanics/ Estimation Theory Symposium on May 16-18, 1995. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Mars Sample Return - Launch and Detection Strategies for Orbital Rendezvous

NASA Technical Reports Server (NTRS)

Woolley, Ryan C.; Mattingly, Richard L.; Riedel, Joseph E.; Sturm, Erick J.

2011-01-01

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/cache rover in 2018, an orbiter with an Earth return vehicle in 2022, and a fetch rover and ascent vehicle in 2024. Strategies are presented to launch the sample into a coplanar orbit with the Orbiter which facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits exist at 457 and 572 km which provide multiple launch opportunities with similar geometries for detection and rendezvous.

Mars Sample Return: Launch and Detection Strategies for Orbital Rendezvous

NASA Technical Reports Server (NTRS)

Woolley, Ryan C.; Mattingly, Richard L.; Riedel, Joseph E.; Sturm, Erick J.

2011-01-01

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/ caching rover in 2018, an Earth return orbiter in 2022, and a fetch rover with ascent vehicle in 2024. Strategies are presented to launch the sample into a nearly coplanar orbit with the Orbiter which would facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits existat 457 and 572 km which would provide multiple launch opportunities with similar geometries for detection and rendezvous.

Rosetta mission operations for landing

NASA Astrophysics Data System (ADS)

Accomazzo, Andrea; Lodiot, Sylvain; Companys, Vicente

2016-08-01

The International Rosetta Mission of the European Space Agency (ESA) was launched on 2nd March 2004 on its 10 year journey to comet Churyumov-Gerasimenko and has reached it early August 2014. The main mission objectives were to perform close observations of the comet nucleus throughout its orbit around the Sun and deliver the lander Philae to its surface. This paper describers the activities at mission operations level that allowed the landing of Philae. The landing preparation phase was mainly characterised by the definition of the landing selection process, to which several parties contributed, and by the definition of the strategy for comet characterisation, the orbital strategy for lander delivery, and the definition and validation of the operations timeline. The definition of the landing site selection process involved almost all components of the mission team; Rosetta has been the first, and so far only mission, that could not rely on data collected by previous missions for the landing site selection. This forced the teams to include an intensive observation campaign as a mandatory part of the process; several science teams actively contributed to this campaign thus making results from science observations part of the mandatory operational products. The time allocated to the comet characterisation phase was in the order of a few weeks and all the processes, tools, and interfaces required an extensive planning an validation. Being the descent of Philae purely ballistic, the main driver for the orbital strategy was the capability to accurately control the position and velocity of Rosetta at Philae's separation. The resulting operations timeline had to merge this need of frequent orbit determination and control with the complexity of the ground segment and the inherent risk of problems when doing critical activities in short times. This paper describes the contribution of the Mission Control Centre (MOC) at the European Space Operations Centre (ESOC) to this mission phase and the lessons learned that can be derived from this experience.

Space Station Freedom altitude strategy

NASA Technical Reports Server (NTRS)

Mcdonald, Brian M.; Teplitz, Scott B.

1990-01-01

The Space Station Freedom (SSF) altitude strategy provides guidelines and assumptions to determine an altitude profile for Freedom. The process for determining an altitude profile incorporates several factors such as where the Space Shuttle will rendezvous with the SSF, when reboosts must occur, and what atmospheric conditions exist causing decay. The altitude strategy has an influence on all areas of SSF development and mission planning. The altitude strategy directly affects the micro-gravity environment for experiments, propulsion and control system sizing, and Space Shuttle delivery manifests. Indirectly the altitude strategy influences almost every system and operation within the Space Station Program. Evolution of the SSF altitude strategy has been a very dynamic process over the past few years. Each altitude strategy in turn has emphasized a different consideration. Examples include a constant Space Shuttle rendezvous altitude for mission planning simplicity, or constant micro-gravity levels with its inherent emphasis on payloads, or lifetime altitudes to provide a safety buffer to loss of control conditions. Currently a new altitude strategy is in development. This altitude strategy will emphasize Space Shuttle delivery optimization. Since propellant is counted against Space Shuttle payload-to-orbit capacity, lowering the rendezvous altitude will not always increase the net payload-to-orbit, since more propellant would be required for reboost. This altitude strategy will also consider altitude biases to account for Space Shuttle launch slips and an unexpected worsening of atmospheric conditions. Safety concerns will define a lower operational altitude limit, while radiation levels will define upper altitude constraints. The evolution of past and current SSF altitude strategies and the development of a new altitude strategy which focuses on operational issues as opposed to design are discussed.

Formations Near the Libration Points: Design Strategies Using Natural and Non-Natural Arcs

NASA Technical Reports Server (NTRS)

Howell, K. C.; Marchand, B. G.

2004-01-01

Space based observatory and interferometry missions, such as Terrestrial Planet Finder (TPF), Stellar Imager, and MAXIM, have sparked great interest in multi-spacecraft formation flight in the vicinity of the Sun-Earth/Moon (SEM) libration points. The initial phase of this research considered the formation keeping problem from the perspective of continuous control as applied to non-natural formations. In the present study, closer inspection of the flow, corresponding to the stable and center manifolds near the reference orbit, reveals some interesting natural relative motions as well as some discrete control strategies for deployment. A hybrid control strategy is also employed that combines both the natural formation dynamics with non-natural motions via input feedback linearization techniques.

Flight Mechanics/Estimation Theory Symposium 1988

NASA Technical Reports Server (NTRS)

Stengle, Thomas (Editor)

1988-01-01

This conference publication includes 28 papers and abstracts presented at the Flight Mechanics/Estimation Theory Symposium on May 10 to 11, 1988. Sponsored by the Flight Dynamics Division of Goddard Space Flight Center, this symposium features technical papers on a wide range of issue related to orbit-attitude prediction, determination and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Solving fuel-optimal low-thrust orbital transfers with bang-bang control using a novel continuation technique

NASA Astrophysics Data System (ADS)

Zhu, Zhengfan; Gan, Qingbo; Yang, Xin; Gao, Yang

2017-08-01

We have developed a novel continuation technique to solve optimal bang-bang control for low-thrust orbital transfers considering the first-order necessary optimality conditions derived from Lawden's primer vector theory. Continuation on the thrust amplitude is mainly described in this paper. Firstly, a finite-thrust transfer with an ;On-Off-On; thrusting sequence is modeled using a two-impulse transfer as initial solution, and then the thrust amplitude is decreased gradually to find an optimal solution with minimum thrust. Secondly, the thrust amplitude is continued from its minimum value to positive infinity to find the optimal bang-bang control, and a thrust switching principle is employed to determine the control structure by monitoring the variation of the switching function. In the continuation process, a bifurcation of bang-bang control is revealed and the concept of critical thrust is proposed to illustrate this phenomenon. The same thrust switching principle is also applicable to the continuation on other parameters, such as transfer time, orbital phase angle, etc. By this continuation technique, fuel-optimal orbital transfers with variable mission parameters can be found via an automated algorithm, and there is no need to provide an initial guess for the costate variables. Moreover, continuation is implemented in the solution space of bang-bang control that is either optimal or non-optimal, which shows that a desired solution of bang-bang control is obtained via continuation on a single parameter starting from an existing solution of bang-bang control. Finally, numerical examples are presented to demonstrate the effectiveness of the proposed continuation technique. Specifically, this continuation technique provides an approach to find multiple solutions satisfying the first-order necessary optimality conditions to the same orbital transfer problem, and a continuation strategy is presented as a preliminary approach for solving the bang-bang control of many-revolution orbital transfers.

Theoretical study and control optimization of an integrated pest management predator-prey model with power growth rate.

PubMed

Sun, Kaibiao; Zhang, Tonghua; Tian, Yuan

2016-09-01

This work presents a pest control predator-prey model, where rate of change in prey density follows a scaling law with exponent less than one and the control is by an integrated management strategy. The aim is to investigate the change in system dynamics and determine a pest control level with minimum control price. First, the dynamics of the proposed model without control is investigated by taking the exponent as an index parameter. And then, to determine the frequency of spraying chemical pesticide and yield releases of the predator, the existence of the order-1 periodic orbit of the control system is discussed in cases. Furthermore, to ensure a certain robustness of the adopted control, i.e., for an inaccurately detected species density or a deviation, the control system could be stabilized at the order-1 periodic orbit, the stability of the order-1 periodic orbit is verified by an stability criterion for a general semi-continuous dynamical system. In addition, to minimize the total cost input in pest control, an optimization problem is formulated and the optimum pest control level is obtained. At last, the numerical simulations with a specific model are carried out to complement the theoretical results. Copyright © 2016 Elsevier Inc. All rights reserved.

Environmental Control and Life Support Integration Strategy for 6-Crew Operations Stephanie Duchesne

NASA Technical Reports Server (NTRS)

Duchesne, Stephanie M.

2009-01-01

The International Space Station (ISS) crew compliment has increased in size from 3 to 6 crew members . In order to support this increase in crew on ISS, the United States on-orbit Segment (USOS) has been outfitted with a suite of regenerative Environmental Control and Life Support (ECLS) hardware including an Oxygen Generation System(OGS), Waste and Hygiene Compartment (WHC), and a Water Recovery System (WRS). The WRS includes the Urine Processor Assembly (UPA) and the Water Processor Assembly (WPA). With this additional life support hardware, the ISS has achieved full redundancy in its on-orbit life support system between the USOS and Russian Segment (RS). The additional redundancy created by the Regenerative ECLS hardware creates the opportunity for independent support capabilities between segments, and for the first time since the start of ISS, the necessity to revise Life Support strategy agreements. Independent operating strategies coupled with the loss of the Space Shuttle supply and return capabilities in 2010 offer new and unique challenges. This paper will discuss the evolution of the ISS Life Support hardware strategy in support of 6-Crew on ISS, as well as the continued work that is necessary to ensure the support of crew and ISS Program objectives through the life of station.

Environmental Control and Life Support Integration Strategy for 6-Crew Operations

NASA Technical Reports Server (NTRS)

Duchesne, Stephanie M.; Tressler, Chad H.

2010-01-01

The International Space Station (ISS) crew complement has increased in size from 3 to 6 crew members. In order to support this increase in crew on ISS, the United States on-orbit Segment (USOS) has been outfitted with a suite of regenerative Environmental Control and Life Support (ECLS) hardware including an Oxygen Generation System (OGS), Waste and Hygiene Compartment (WHC), and a Water Recovery System (WRS). The WRS includes the Urine Processor Assembly (UPA) and the Water Processor Assembly (WPA). With this additional life support hardware, the ISS has achieved full redundancy in its on-orbit life support system between the t OS and Russian Segment (RS). The additional redundancy created by the Regenerative ECLS hardware creates the opportunity for independent support capabilities between segments, and for the first time since the start of ISS, the necessity to revise Life Support strategy agreements. Independent operating strategies coupled with the loss of the Space Shuttle supply and return capabilities in 2010 offer new and unique challenges. This paper will discuss the evolution of the ISS Life Support hardware strategy in support of 6-Crew on ISS, as well as the continued work that is necessary to ensure the support of crew and ISS Program objectives through the life of station

Orbital Injection of the SEDSAT Satellite: Tethered Systems Dynamics and Flight Data Analysis

NASA Technical Reports Server (NTRS)

Lorenzini, Enrico C.; Gullahorn, Gordon E.; Cosmo, Mario L.; Ruiz, Manuel; Pelaez, Jesus

1996-01-01

This report deals with the following topics which are all related to the orbital injection of the SEDSAT satellite: Dynamics and Stability of Tether Oscillations after the First Cut. The dynamics of the tether after the first cut (i.e., without the Shuttle attached to it) is investigated. The tether oscillations with the free end are analyzed in order to assess the stability of the rectilinear configuration in between the two tether cuts; analysis of Unstable Modes. The unstable modes that appear for high libration angles are further investigated in order to determine their occurrences and the possible transition from bound librations to rotations; Orbital Release Strategies for SEDSAT. A parametric analysis of the orbital decay rate of the SEDSAT satellite after the two tether cuts has been carried out as a function of the following free parameters: libration amplitude at the end of deployment, deviation angle from LV at the first cut, and orbital anomaly at the second cut. The values of these parameters that provide a minimum orbital decay rate of the satellite (after the two cuts) have been computed; and Dynamics and Control of SEDSAT. The deployment control law has been modified to cope with the new ejection velocity of the satellite from the Shuttle cargo bay. New reference profiles have been derived as well as new control parameters. Timing errors at the satellite release as a function of the variations of the initial conditions and the tension model parameters have been estimated for the modified control law.

Integrated guidance and control for microsatellite real-time automated proximity operations

NASA Astrophysics Data System (ADS)

Chen, Ying; He, Zhen; Zhou, Ding; Yu, Zhenhua; Li, Shunli

2018-07-01

This paper investigates the trajectory planning and control of autonomous spacecraft proximity operations with impulsive dynamics. A new integrated guidance and control scheme is developed to perform automated close-range rendezvous for underactuated microsatellites. To efficiently prevent collision, a modified RRT* trajectory planning algorithm is proposed under this context. Several engineering constraints such as collision avoidance, plume impingement, field of view and control feasibility are considered simultaneously. Then, the feedback controller that employs a turn-burn-turn strategy with a combined impulsive orbital control and finite-time attitude control is designed to ensure the implementation of planned trajectory. Finally, the performance of trajectory planner and controller are evaluated through numerical tests. Simulation results indicate the real-time implementability of the proposed integrated guidance and control scheme with position control error less than 0.5 m and velocity control error less than 0.05 m/s. Consequently, the proposed scheme offers the potential for wide applications, such as on-orbit maintenance, space surveillance and debris removal.

Frozen orbit realization using LQR analogy

NASA Astrophysics Data System (ADS)

Nagarajan, N.; Rayan, H. Reno

In the case of remote sensing orbits, the Frozen Orbit concept minimizes altitude variations over a given region using passive means. This is achieved by establishing the mean eccentricity vector at the orbital poles i.e., by fixing the mean argument of perigee at 90 deg with an appropriate eccentricity to balance the perturbations due to zonal harmonics J2 and J3 of the Earth's potential. Eccentricity vector is a vector whose magnitude is the eccentricity and direction is the argument of perigee. The launcher dispersions result in an eccentricity vector which is away from the frozen orbit values. The objective is then to formulate an orbit maneuver strategy to optimize the fuel required to achieve the frozen orbit in the presence of visibility and impulse constraints. It is shown that the motion of the eccentricity vector around the frozen perigee can be approximated as a circle. Combining the circular motion of the eccentricity vector around the frozen point and the maneuver equation, the following discrete equation is obtained. X(k+1) = AX(k) + Bu(k), where X is the state (i.e. eccentricity vector components), A the state transition matrix, u the scalar control force (i.e. dV in this case) and B the control matrix which transforms dV into eccentricity vector change. Based on this, it is shown that the problem of optimizing the fuel can be treated as a Linear Quadratic Regulator (LQR) problem in which the maneuver can be solved by using control system design tools like MATLAB by deriving an analogy LQR design.

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

NASA Astrophysics Data System (ADS)

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 with implementation of electrostatic detumble control on-orbit. Leveraging an extended Kalman filter scheme, the relative position information is readily obtained. In order to touchlessly acquire the target electrostatic potential, a nested two-time scale Kalman filter is employed to provide real-time estimates of both relative position and electrostatic potential while on-orbit. The culmination of the presented control formulations for generic spacecraft geometries, the proximity and formation flying control capability, and the availability of necessary state information provide significant contributions towards the viability of electrostatic detumble mission concepts.

A Survey of Ballistic Transfers to Low Lunar Orbit

NASA Technical Reports Server (NTRS)

Parker, Jeffrey S.; Anderson, Rodney L.; Peterson, Andrew

2011-01-01

A simple strategy is identified to generate ballistic transfers between the Earth and Moon, i.e., transfers that perform two maneuvers: a trans-lunar injection maneuver to depart the Earth and a Lunar Orbit Insertion maneuver to insert into orbit at the Moon. This strategy is used to survey the performance of numerous transfers between varying Earth parking orbits and varying low lunar target orbits. The transfers surveyed include short 3-6 day direct transfers, longer 3-4 month low energy transfers, and variants that include Earth phasing orbits and/or lunar flybys.

A Minimum (Delta)V Orbit Maintenance Strategy for Low-Altitude Missions Using Burn Parameter Optimization

NASA Technical Reports Server (NTRS)

Brown, Aaron J.

2011-01-01

Orbit maintenance is the series of burns performed during a mission to ensure the orbit satisfies mission constraints. Low-altitude missions often require non-trivial orbit maintenance (Delta)V due to sizable orbital perturbations and minimum altitude thresholds. A strategy is presented for minimizing this (Delta)V using impulsive burn parameter optimization. An initial estimate for the burn parameters is generated by considering a feasible solution to the orbit maintenance problem. An example demonstrates the dV savings from the feasible solution to the optimal solution.

Bi-objective optimization of a multiple-target active debris removal mission

NASA Astrophysics Data System (ADS)

Bérend, Nicolas; Olive, Xavier

2016-05-01

The increasing number of space debris in Low-Earth Orbit (LEO) raises the question of future Active Debris Removal (ADR) operations. Typical ADR scenarios rely on an Orbital Transfer Vehicle (OTV) using one of the two following disposal strategies: the first one consists in attaching a deorbiting kit, such as a solid rocket booster, to the debris after rendezvous; with the second one, the OTV captures the debris and moves it to a low-perigee disposal orbit. For multiple-target ADR scenarios, the design of such a mission is very complex, as it involves two optimization levels: one for the space debris sequence, and a second one for the "elementary" orbit transfer strategy from a released debris to the next one in the sequence. This problem can be seen as a Time-Dependant Traveling Salesman Problem (TDTSP) with two objective functions to minimize: the total mission duration and the total propellant consumption. In order to efficiently solve this problem, ONERA has designed, under CNES contract, TOPAS (Tool for Optimal Planning of ADR Sequence), a tool that implements a Branch & Bound method developed in previous work together with a dedicated algorithm for optimizing the "elementary" orbit transfer. A single run of this tool yields an estimation of the Pareto front of the problem, which exhibits the trade-off between mission duration and propellant consumption. We first detail our solution to cope with the combinatorial explosion of complex ADR scenarios with 10 debris. The key point of this approach is to define the orbit transfer strategy through a small set of parameters, allowing an acceptable compromise between the quality of the optimum solution and the calculation cost. Then we present optimization results obtained for various 10 debris removal scenarios involving a 15-ton OTV, using either the deorbiting kit or the disposal orbit strategy. We show that the advantage of one strategy upon the other depends on the propellant margin, the maximum duration allowed for the mission and the orbit inclination domain. For high inclination orbits near 98 deg, the disposal orbit strategy is more appropriate for short duration missions, while the deorbiting kit strategy ensures a better propellant margin. Conversely, for lower inclination orbits near 65 deg, the deorbiting kit strategy appears to be the only possible with a 10 debris set. We eventually explain the consistency of these results with regards to astrodynamics.

Shared control on lunar spacecraft teleoperation rendezvous operations with large time delay

NASA Astrophysics Data System (ADS)

Ya-kun, Zhang; Hai-yang, Li; Rui-xue, Huang; Jiang-hui, Liu

2017-08-01

Teleoperation could be used in space on-orbit serving missions, such as object deorbits, spacecraft approaches, and automatic rendezvous and docking back-up systems. Teleoperation rendezvous and docking in lunar orbit may encounter bottlenecks for the inherent time delay in the communication link and the limited measurement accuracy of sensors. Moreover, human intervention is unsuitable in view of the partial communication coverage problem. To solve these problems, a shared control strategy for teleoperation rendezvous and docking is detailed. The control authority in lunar orbital maneuvers that involves two spacecraft as rendezvous and docking in the final phase was discussed in this paper. The predictive display model based on the relative dynamic equations is established to overcome the influence of the large time delay in communication link. We discuss and attempt to prove via consistent, ground-based simulations the relative merits of fully autonomous control mode (i.e., onboard computer-based), fully manual control (i.e., human-driven at the ground station) and shared control mode. The simulation experiments were conducted on the nine-degrees-of-freedom teleoperation rendezvous and docking simulation platform. Simulation results indicated that the shared control methods can overcome the influence of time delay effects. In addition, the docking success probability of shared control method was enhanced compared with automatic and manual modes.

Stationkeeping of Lissajous Trajectories in the Earth-Moon System with Applications to ARTEMIS

NASA Technical Reports Server (NTRS)

Folta, D. C.; Pavlak, T. A.; Howell, K. C.; Woodard, M. A.; Woodfork, D. W.

2010-01-01

In the last few decades, several missions have successfully exploited trajectories near the.Sun-Earth L1 and L2 libration points. Recently, the collinear libration points in the Earth-Moon system have emerged as locations with immediate application. Most libration point orbits, in any system, are inherently unstable. and must be controlled. To this end, several stationkeeping strategies are considered for application to ARTEMIS. Two approaches are examined to investigate the stationkeeping problem in this regime and the specific options. available for ARTEMIS given the mission and vehicle constraints. (I) A baseline orbit-targeting approach controls the vehicle to remain near a nominal trajectory; a related global optimum search method searches all possible maneuver angles to determine an optimal angle and magnitude; and (2) an orbit continuation method, with various formulations determines maneuver locations and minimizes costs. Initial results indicate that consistent stationkeeping costs can be achieved with both approaches and the costs are reasonable. These methods are then applied to Lissajous trajectories representing a baseline ARTEMIS libration orbit trajectory.

Orbit determination strategy and results for the Pioneer 10 Jupiter mission

NASA Technical Reports Server (NTRS)

Wong, S. K.; Lubeley, A. J.

1974-01-01

Pioneer 10 is the first earth-based vehicle to encounter Jupiter and occult its moon, Io. In contributing to the success of the mission, the Orbit Determination Group evaluated the effects of the dominant error sources on the spacecraft's computed orbit and devised an encounter strategy minimizing the effects of these error sources. The encounter results indicated that: (1) errors in the satellite model played a very important role in the accuracy of the computed orbit, (2) encounter strategy was sound, (3) all mission objectives were met, and (4) Jupiter-Saturn mission for Pioneer 11 is within the navigation capability.

The dynamics and control of large-flexible space structures, part 10

NASA Technical Reports Server (NTRS)

Bainum, Peter M.; Reddy, A. S. S. R.

1988-01-01

A mathematical model is developed to predict the dynamics of the proposed orbiting Spacecraft Control Laboratory Experiment (SCOLE) during the station keeping phase. The equations of motion are derived using a Newton-Euler formulation. The model includes the effects of gravity, flexibility, and orbital dynamics. The control is assumed to be provided to the system through the Shuttle's three torquers, and through six actuators located by pairs at two points on the mast and at the mass center of the reflector. The modal shape functions are derived using the fourth order beam equation. The generic mode equations are derived to account for the effects of the control forces on the modal shape and frequencies. The equations are linearized about a nominal equilibrium position. The linear regulator theory is used to derive control laws for both the linear model of the rigidized SCOLE as well as that of the actual SCOLE including the first four flexible modes. The control strategy previously derived for the linear model of the rigidized SCOLE is applied to the nonlinear model of the same configuration of the system and preliminary single axis slewing maneuvers conducted. The results obtained confirm the applicability of the intuitive and appealing two-stage control strategy which would slew the SCOLE system, as if rigid to its desired position and then concentrate on damping out the residual flexible motions.

Development of a Two-Wheel Contingency Mode for the MAP Spacecraft

NASA Technical Reports Server (NTRS)

Starin, Scott R.; ODonnell, James R., Jr.; Bauer, Frank (Technical Monitor)

2002-01-01

The Microwave Anisotropy Probe (MAP) is a follow-on mission to the Cosmic Background Explorer (COBE), and is currently collecting data from its orbit near the second Sun-Earth libration point. Due to limited mass, power, and financial resources, a traditional reliability concept including fully redundant components was not feasible for MAP. Instead, the MAP design employs selective hardware redundancy in tandem with contingency software modes and algorithms to improve the odds of mission success. One direction for such improvement has been the development of a two-wheel backup control strategy. This strategy would allow MAP to position itself for maneuvers and collect science data should one of its three reaction wheels fail. Along with operational considerations, the strategy includes three new control algorithms. These algorithms would use the remaining attitude control actuators-thrusters and two reaction wheels-in ways that achieve control goals while minimizing adverse impacts on the functionality of other subsystems and software.

Precise Trajectory Reconstruction of CE-3 Hovering Stage By Landing Camera Images

NASA Astrophysics Data System (ADS)

Yan, W.; Liu, J.; Li, C.; Ren, X.; Mu, L.; Gao, X.; Zeng, X.

2014-12-01

Chang'E-3 (CE-3) is part of the second phase of the Chinese Lunar Exploration Program, incorporating a lander and China's first lunar rover. It was landed on 14 December, 2013 successfully. Hovering and obstacle avoidance stages are essential for CE-3 safety soft landing so that precise spacecraft trajectory in these stages are of great significance to verify orbital control strategy, to optimize orbital design, to accurately determine the landing site of CE-3, and to analyze the geological background of the landing site. Because the time consumption of these stages is just 25s, it is difficult to present spacecraft's subtle movement by Measurement and Control System or by radio observations. Under this background, the trajectory reconstruction based on landing camera images can be used to obtain the trajectory of CE-3 because of its technical advantages such as unaffecting by lunar gravity field spacecraft kinetic model, high resolution, high frame rate, and so on. In this paper, the trajectory of CE-3 before and after entering hovering stage was reconstructed by landing camera images from frame 3092 to frame 3180, which lasted about 9s, under Single Image Space Resection (SISR). The results show that CE-3's subtle changes during hovering stage can be emerged by the reconstructed trajectory. The horizontal accuracy of spacecraft position was up to 1.4m while vertical accuracy was up to 0.76m. The results can be used for orbital control strategy analysis and some other application fields.

Guidance of a Solar Sail Spacecraft to the Sun - L(2) Point.

NASA Astrophysics Data System (ADS)

Hur, Sun Hae

The guidance of a solar sail spacecraft along a minimum-time path from an Earth orbit to a region near the Sun-Earth L_2 libration point is investigated. Possible missions to this point include a spacecraft "listening" for possible extra-terrestrial electromagnetic signals and a science payload to study the geomagnetic tail. A key advantage of the solar sail is that it requires no fuel. The control variables are the sail angles relative to the Sun-Earth line. The thrust is very small, on the order of 1 mm/s^2, and its magnitude and direction are highly coupled. Despite this limited controllability, the "free" thrust can be used for a wide variety of terminal conditions including halo orbits. If the Moon's mass is lumped with the Earth, there are quasi-equilibrium points near L_2. However, they are unstable so that some form of station keeping is required, and the sail can provide this without any fuel usage. In the two-dimensional case, regulating about a nominal orbit is shown to require less control and result in smaller amplitude error response than regulating about a quasi-equilibrium point. In the three-dimensional halo orbit case, station keeping using periodically varying gains is demonstrated. To compute the minimum-time path, the trajectory is divided into two segments: the spiral segment and the transition segment. The spiral segment is computed using a control law that maximizes the rate of energy increase at each time. The transition segment is computed as the solution of the time-optimal control problem from the endpoint of the spiral to the terminal point. It is shown that the path resulting from this approximate strategy is very close to the exact optimal path. For the guidance problem, the approximate strategy in the spiral segment already gives a nonlinear full-state feedback law. However, for large perturbations, follower guidance using an auxiliary propulsion is used for part of the spiral. In the transition segment, neighboring extremal feedback guidance using the solar sail, with feedforward control only near the terminal point, is used to correct perturbations in the initial conditions.

Human Exploration using Real-Time Robotic Operations (HERRO): A space exploration strategy for the 21st century

NASA Astrophysics Data System (ADS)

Schmidt, George R.; Landis, Geoffrey A.; Oleson, Steven R.

2012-11-01

This paper presents an exploration strategy for human missions beyond Low Earth Orbit (LEO) and the Moon that combines the best features of human and robotic spaceflight. This "Human Exploration using Real-time Robotic Operations" (HERRO) strategy refrains from placing humans on the surfaces of the Moon and Mars in the near-term. Rather, it focuses on sending piloted spacecraft and crews into orbit around Mars and other exploration targets of interest, and conducting astronaut exploration of the surfaces using telerobots and remotely-controlled systems. By eliminating the significant communications delay or "latency" with Earth due to the speed of light limit, teleoperation provides scientists real-time control of rovers and other sophisticated instruments. This in effect gives them a "virtual presence" on planetary surfaces, and thus expands the scientific return at these destinations. HERRO mitigates several of the major issues that have hindered the progress of human spaceflight beyond Low Earth Orbit (LEO) by: (1) broadening the range of destinations for near-term human missions; (2) reducing cost and risk through less complexity and fewer man-rated elements; (3) offering benefits of human-equivalent in-situ cognition, decision-making and field-work on planetary bodies; (4) providing a simpler approach to returning samples from Mars and planetary surfaces; and (5) facilitating opportunities for international collaboration through contribution of diverse robotic systems. HERRO provides a firm justification for human spaceflight—one that expands the near-term capabilities of scientific exploration while providing the space transportation infrastructure needed for eventual human landings in the future.

Tether fundamentals

NASA Technical Reports Server (NTRS)

Carroll, J. A.

1986-01-01

Some fundamental aspects of tethers are presented and briefly discussed. The effects of gravity gradients, dumbbell libration in circular orbits, tether control strategies and impact hazards for tethers are among those fundamentals. Also considered are aerodynamic drag, constraints in momentum transfer applications and constraints with permanently deployed tethers. The theoretical feasibility of these concepts are reviewed.

Fast Low-Current Spin-Orbit-Torque Switching of Magnetic Tunnel Junctions through Atomic Modifications of the Free-Layer Interfaces

NASA Astrophysics Data System (ADS)

Shi, Shengjie; Ou, Yongxi; Aradhya, S. V.; Ralph, D. C.; Buhrman, R. A.

2018-01-01

Future applications of spin-orbit torque will require new mechanisms to improve the efficiency of switching nanoscale magnetic tunnel junctions (MTJs), while also controlling the magnetic dynamics to achieve fast nanosecond-scale performance with low-write-error rates. Here, we demonstrate a strategy to simultaneously enhance the interfacial magnetic anisotropy energy and suppress interfacial spin-memory loss by introducing subatomic and monatomic layers of Hf at the top and bottom interfaces of the ferromagnetic free layer of an in-plane magnetized three-terminal MTJ device. When combined with a β -W spin Hall channel that generates spin-orbit torque, the cumulative effect is a switching current density of 5.4 ×106 A /cm2 .

TOPEX/POSEIDON orbit maintenance maneuver design

NASA Technical Reports Server (NTRS)

Bhat, R. S.; Frauenholz, R. B.; Cannell, Patrick E.

1990-01-01

The Ocean Topography Experiment (TOPEX/POSEIDON) mission orbit requirements are outlined, as well as its control and maneuver spacing requirements including longitude and time targeting. A ground-track prediction model dealing with geopotential, luni-solar gravity, and atmospheric-drag perturbations is considered. Targeting with all modeled perturbations is discussed, and such ground-track prediction errors as initial semimajor axis, orbit-determination, maneuver-execution, and atmospheric-density modeling errors are assessed. A longitude targeting strategy for two extreme situations is investigated employing all modeled perturbations and prediction errors. It is concluded that atmospheric-drag modeling errors are the prevailing ground-track prediction error source early in the mission during high solar flux, and that low solar-flux levels expected late in the experiment stipulate smaller maneuver magnitudes.

Control of spin-orbit torques through crystal symmetry in WTe2/ferromagnet bilayers

NASA Astrophysics Data System (ADS)

MacNeill, D.; Stiehl, G. M.; Guimaraes, M. H. D.; Buhrman, R. A.; Park, J.; Ralph, D. C.

2017-03-01

Recent discoveries regarding current-induced spin-orbit torques produced by heavy-metal/ferromagnet and topological-insulator/ferromagnet bilayers provide the potential for dramatically improved efficiency in the manipulation of magnetic devices. However, in experiments performed to date, spin-orbit torques have an important limitation--the component of torque that can compensate magnetic damping is required by symmetry to lie within the device plane. This means that spin-orbit torques can drive the most current-efficient type of magnetic reversal (antidamping switching) only for magnetic devices with in-plane anisotropy, not the devices with perpendicular magnetic anisotropy that are needed for high-density applications. Here we show experimentally that this state of affairs is not fundamental, but rather one can change the allowed symmetries of spin-orbit torques in spin-source/ferromagnet bilayer devices by using a spin-source material with low crystalline symmetry. We use WTe2, a transition-metal dichalcogenide whose surface crystal structure has only one mirror plane and no two-fold rotational invariance. Consistent with these symmetries, we generate an out-of-plane antidamping torque when current is applied along a low-symmetry axis of WTe2/Permalloy bilayers, but not when current is applied along a high-symmetry axis. Controlling spin-orbit torques by crystal symmetries in multilayer samples provides a new strategy for optimizing future magnetic technologies.

On-board orbit determination for low thrust LEO-MEO transfer by Consider Kalman Filtering and multi-constellation GNSS

NASA Astrophysics Data System (ADS)

Menzione, Francesco; Renga, Alfredo; Grassi, Michele

2017-09-01

In the framework of the novel navigation scenario offered by the next generation satellite low thrust autonomous LEO-to-MEO orbit transfer, this study proposes and tests a GNSS based navigation system aimed at providing on-board precise and robust orbit determination strategy to override rising criticalities. The analysis introduces the challenging design issues to simultaneously deal with the variable orbit regime, the electric thrust control and the high orbit GNSS visibility conditions. The Consider Kalman Filtering approach is here proposed as the filtering scheme to process the GNSS raw data provided by a multi-antenna/multi-constellation receiver in presence of uncertain parameters affecting measurements, actuation and spacecraft physical properties. Filter robustness and achievable navigation accuracy are verified using a high fidelity simulation of the low-thrust rising scenario and performance are compared with the one of a standard Extended Kalman Filtering approach to highlight the advantages of the proposed solution. Performance assessment of the developed navigation solution is accomplished for different transfer phases.

Patient-specific puzzle implant preformed with 3D-printed rapid prototype model for combined orbital floor and medial wall fracture.

PubMed

Kim, Young Chul; Min, Kyung Hyun; Choi, Jong Woo; Koh, Kyung S; Oh, Tae Suk; Jeong, Woo Shik

2018-04-01

The management of combined orbital floor and medial wall fractures involving the inferomedial strut is challenging due to absence of stable cornerstone. In this article, we proposed surgical strategies using customized 3D puzzle implant preformed with Rapid Prototype (RP) skull model. Retrospective review was done in 28 patients diagnosed with combined orbital floor and medial wall fracture. Using preoperative CT scans, original and mirror-imaged RP skull models for each patient were prepared and sterilized. In all patients, porous polyethylene-coated titanium mesh was premolded onto RP skull model in two ways; Customized 3D jigsaw puzzle technique was used in 15 patients with comminuted inferomedial strut, whereas individual 3D implant technique was used in each fracture for 13 patients with intact inferomedial strut. Outcomes including enophthalmos, visual acuity, and presence of diplopia were assessed and orbital volume was measured using OsiriX software preoperatively and postoperatively. Satisfactory results were achieved in both groups in terms of clinical improvements. Of 10 patients with preoperative diplopia, 9 improved in 6 months, except one with persistent symptom who underwent extraocular muscle rupture. 18 patients who had moderate to severe enophthalmos preoperatively improved, and one remained with mild degree. Orbital volume ratio, defined as volumetric ratio between affected and control orbit, decreased from 127.6% to 99.79% (p < 0.05) in comminuted group, and that in intact group decreased from 117.03% to 101.3% (p < 0.05). Our surgical strategies using the jigsaw puzzle and individual reconstruction technique provide accurate restoration of combined orbital floor and medial wall fractures. Copyright © 2017 British Association of Plastic, Reconstructive and Aesthetic Surgeons. Published by Elsevier Ltd. All rights reserved.

Fuel-optimal trajectories of aeroassisted orbital transfer with plane change

NASA Technical Reports Server (NTRS)

Naidu, Desineni Subbaramaiah; Hibey, Joseph L.

1989-01-01

The problem of minimization of fuel consumption during the atmospheric portion of an aeroassisted, orbital transfer with plane change is addressed. The complete mission has required three characteristic velocities, a deorbit impulse at high earth orbit (HEO), a boost impulse at the atmospheric exit, and a reorbit impulse at low earth orbit (LEO). A performance index has been formulated as the sum of these three impulses. Application of optimal control principles has led to a nonlinear, two-point, boundary value problem which was solved by using a multiple shooting algorithm. The strategy for the atmospheric portion of the minimum-fuel transfer is to start initially with the maximum positive lift in order to recover from the downward plunge, and then to fly with a gradually decreasing lift such that the vehicle skips out of the atmosphere with a flight path angle near zero degrees.

Design and implementation of satellite formations and constellations

NASA Technical Reports Server (NTRS)

Folta, David; Newman, Lauri Kraft; Quinn, David

1998-01-01

The direction to develop small low cost spacecraft has led many scientists to recognize the advantage of flying spacecraft in constellations and formations to achieve the correlated instrument measurements formerly possible only by flying many instruments on a single large platform. Yet, constellations and formation flying impose additional complications on orbit selection and orbit maintenance, especially when each spacecraft has its own orbit or science requirements. The purpose of this paper is to develop an operational control method for maintenance of these missions. Examples will be taken from the Earth Observing-1 (EO-1) spacecraft that is part of the New Millennium Program (NMP) and from proposed Earth System Science Program Office (ESSPO) constellations. Results can be used to determine the appropriateness of constellations and formation flying for a particular case as well as the operational impacts. Applications to the ESSPO and NMP are highly considered in analysis and applications. After constellation and formation analysis is completed, implementation of a maneuver maintenance strategy becomes the driver. Advances in technology and automation by GSFC's Guidance, Navigation, and Control Center allow more of the burden of the orbit selection and maneuver maintenance to be automated and ultimately placed onboard the spacecraft, mitigating most of the associated operational concerns. This paper presents the GSFC closed-loop control method to fly in either constellations or formations through the use of an autonomous closed loop three-axis navigation control and innovative orbit maintenance support. Simulation results using AutoCon(TM) and FreeFlyer(TM) with various fidelity levels of modeling and algorithms are presented.

Design and Implementation of Satellite Formations and Constellations

NASA Technical Reports Server (NTRS)

Folta, David; Newman, Lauri Kraft; Quinn, David

1998-01-01

The direction to develop small low cost spacecraft has led many scientists to recognize the advantage of flying spacecraft in constellations and formations to achieve the correlated instrument measurements formerly possible only by flying many instruments on a single large platform. Yet, constellations and formation flying impose additional complications on orbit selection and orbit maintenance, especially when each spacecraft has its own orbit or science requirements. The purpose of this paper is to develop an operational control method for maintenance of these missions. Examples will be taken from the Earth Observing-1 (EO-1) spacecraft that is part of the New Millennium Program (NMP) and from proposed Earth System Science Program Office (ESSPO) constellations. Results can be used to determine the appropriateness of constellations and formation flying for a particular case as well as the operational impacts. Applications to the ESSPO and NMP are highly considered in analysis and applications. After constellation and formation analysis is completed, implementation of a maneuver maintenance strategy becomes the driver. Advances in technology and automation by GSFC's Guidance, Navigation, and Control Center allow more of the burden of the orbit selection and maneuver maintenance to be automated and ultimately placed onboard the spacecraft, mitigating most of the associated operational concerns. This paper presents the GSFC closed-loop control method to fly in either constellations or formations through the use of an autonomous closed loop three-axis navigation control and innovative orbit maintenance support. Simulation results using AutoCon(Trademark) and FreeFlyer(Trademark) with various fidelity levels of modeling and algorithms are presented.

Hardware-In-The-Loop Testing of Continuous Control Algorithms for a Precision Formation Flying Demonstration Mission

NASA Technical Reports Server (NTRS)

Naasz, Bo J.; Burns, Richard D.; Gaylor, David; Higinbotham, John

2004-01-01

A sample mission sequence is defined for a low earth orbit demonstration of Precision Formation Flying (PFF). Various guidance navigation and control strategies are discussed for use in the PFF experiment phases. A sample PFF experiment is implemented and tested in a realistic Hardware-in-the-Loop (HWIL) simulation using the Formation Flying Test Bed (FFTB) at NASA's Goddard Space Flight Center.

Space Shuttle flying qualities criteria assessment. Phase 5: Data acquistion and analysis

NASA Technical Reports Server (NTRS)

Myers, T. T.; Johnston, D. E.; Mcruer, D. T.

1986-01-01

The development of flying qualities experiments (OFQ) as a part of the Orbiter Experiments Program (OEX) was continued. The data base was extended to use the ground based cinetheodolite measurements of orbiter approach and landing. Onboard the cinetheodolite data were analyzed from flights STS 2 through 7 to identify the effective augmented vehicle dynamics, the control strategy employed by the pilot during preflare, shallow glide, and final flare segments of the landing, and the key approach and touchdown performance measures. A plan for an OFQ flying qualities data archive and processing is presented.

Codimension-1 Sliding Bifurcations of a Filippov Pest Growth Model with Threshold Policy

NASA Astrophysics Data System (ADS)

Tang, Sanyi; Tang, Guangyao; Qin, Wenjie

A Filippov system is proposed to describe the stage structured nonsmooth pest growth with threshold policy control (TPC). The TPC measure is represented by the total density of both juveniles and adults being chosen as an index for decisions on when to implement chemical control strategies. The proposed Filippov system can have three pieces of sliding segments and three pseudo-equilibria, which result in rich sliding mode bifurcations and local sliding bifurcations including boundary node (boundary focus, or boundary saddle) and tangency bifurcations. As the threshold density varies the model exhibits the interesting global sliding bifurcations sequentially: touching → buckling → crossing → sliding homoclinic orbit to a pseudo-saddle → crossing → touching bifurcations. In particular, bifurcation of a homoclinic orbit to a pseudo-saddle with a figure of eight shape, to a pseudo-saddle-node or to a standard saddle-node have been observed for some parameter sets. This implies that control outcomes are sensitive to the threshold level, and hence it is crucial to choose the threshold level to initiate control strategy. One more sliding segment (or pseudo-equilibrium) is induced by the total density of a population guided switching policy, compared to only the juvenile density guided policy, implying that this control policy is more effective in terms of preventing multiple pest outbreaks or causing the density of pests to stabilize at a desired level such as an economic threshold.

Filter Strategies for Mars Science Laboratory Orbit Determination

NASA Technical Reports Server (NTRS)

Thompson, Paul F.; Gustafson, Eric D.; Kruizinga, Gerhard L.; Martin-Mur, Tomas J.

2013-01-01

The Mars Science Laboratory (MSL) spacecraft had ambitious navigation delivery and knowledge accuracy requirements for landing inside Gale Crater. Confidence in the orbit determination (OD) solutions was increased by investigating numerous filter strategies for solving the orbit determination problem. We will discuss the strategy for the different types of variations: for example, data types, data weights, solar pressure model covariance, and estimating versus considering model parameters. This process generated a set of plausible OD solutions that were compared to the baseline OD strategy. Even implausible or unrealistic results were helpful in isolating sensitivities in the OD solutions to certain model parameterizations or data types.

Space Weather Influence on Relative Motion Control using the Touchless Electrostatic Tractor

NASA Astrophysics Data System (ADS)

Hogan, Erik A.; Schaub, Hanspeter

2016-09-01

With recent interest in the use of electrostatic forces for contactless tugging and attitude control of noncooperative objects for orbital servicing and active debris mitigation, the need for a method of remote charge control arises. In this paper, the use of a directed electron beam for remote charge control is considered in conjunction with the relative motion control. A tug vehicle emits an electron beam onto a deputy object, charging it negatively. At the same time, the tug is charged positively due to beam emission, resulting in an attractive electrostatic force. The relative position feedback control between the tug and the passive debris object is studied subject to the charging being created through an electron beam. Employing the nominal variations of the GEO space weather conditions across longitude slots, two electrostatic tugging strategies are considered. First, the electron beam current is adjusted throughout the orbit in order to maximize this resulting electrostatic force. This open-loop control strategy compensates for changes in the nominally expected local space weather environment in the GEO region to adjust for fluctuations in the local plasma return currents. Second, the performance impact of using a fixed electron beam current on the electrostatic tractor is studied if the same natural space weather variations are assumed. The fixed electron beam current shows a minor performance penalty (<5 %) while providing a much simpler implementation that does not require any knowledge of local space weather conditions.

Automatic trajectory planning for low-thrust active removal mission in low-earth orbit

NASA Astrophysics Data System (ADS)

Di Carlo, Marilena; Romero Martin, Juan Manuel; Vasile, Massimiliano

2017-03-01

In this paper two strategies are proposed to de-orbit up to 10 non-cooperative objects per year from the region within 800 and 1400 km altitude in Low Earth Orbit (LEO). The underlying idea is to use a single servicing spacecraft to de-orbit several objects applying two different approaches. The first strategy is analogous to the Traveling Salesman Problem: the servicing spacecraft rendezvous with multiple objects in order to physically attach a de-orbiting kit that reduces the perigee of the orbit. The second strategy is analogous to the Vehicle Routing Problem: the servicing spacecraft rendezvous and docks with an object, spirals it down to a lower altitude orbit, undocks, and then spirals up to the next target. In order to maximise the number of de-orbited objects with minimum propellant consumption, an optimal sequence of targets is identified using a bio-inspired incremental automatic planning and scheduling discrete optimisation algorithm. The optimisation of the resulting sequence is realised using a direct transcription method based on an asymptotic analytical solution of the perturbed Keplerian motion. The analytical model takes into account the perturbations deriving from the J2 gravitational effect and the atmospheric drag.

HERRO: A Science-Oriented Strategy for Crewed Missions Beyond LEO

NASA Technical Reports Server (NTRS)

Schmidt, George R.

2011-01-01

This paper presents an exploration strategy for human missions beyond Low Earth Orbit (LEO) and the Moon that combines the best features of human and robotic spaceflight. This "Human Exploration using Real-time Robotic Operations" (HERRO) strategy refrains from placing humans on the surfaces of the Moon and Mars in the near-term. Rather, it focuses on sending piloted spacecraft and crews into orbit around exploration targets of interest, such as Mars, and conducting astronaut exploration of the surfaces using telerobots and remotely controlled systems. By eliminating the significant communications delay with Earth due to the speed of light limit, teleoperation provides scientists real-time control of rovers and other sophisticated instruments, in effect giving them a "virtual presence" on planetary surfaces, and thus expanding the scientific return at these destinations. It also eliminates development of the numerous man-rated landers, ascent vehicles and surface systems that are required to land humans on planetary surfaces. The propulsive requirements to travel from LEO to many destinations with shallow gravity-wells in the inner solar system are quite similar. Thus, a single spacecraft design could perform a variety of missions, including orbit-based surface exploration of the Moon, Mars and Venus, and rendezvous with Near Earth Asteroids (NEAs), as well as Phobos and Deimos. Although HERRO bypasses many of the initial steps that have been historically associated with human space exploration, it opens the door to many new destinations that are candidates for future resource utilization and settlement. HERRO is a first step that takes humans to exciting destinations beyond LEO, while expanding the ability to conduct science within the inner solar system.

Method of resolving radio phase ambiguity in satellite orbit determination

NASA Technical Reports Server (NTRS)

Councelman, Charles C., III; Abbot, Richard I.

1989-01-01

For satellite orbit determination, the most accurate observable available today is microwave radio phase, which can be differenced between observing stations and between satellites to cancel both transmitter- and receiver-related errors. For maximum accuracy, the integer cycle ambiguities of the doubly differenced observations must be resolved. To perform this ambiguity resolution, a bootstrapping strategy is proposed. This strategy requires the tracking stations to have a wide ranging progression of spacings. By conventional 'integrated Doppler' processing of the observations from the most widely spaced stations, the orbits are determined well enough to permit resolution of the ambiguities for the most closely spaced stations. The resolution of these ambiguities reduces the uncertainty of the orbit determination enough to enable ambiguity resolution for more widely spaced stations, which further reduces the orbital uncertainty. In a test of this strategy with six tracking stations, both the formal and the true errors of determining Global Positioning System satellite orbits were reduced by a factor of 2.

Accurate orbit determination strategies for the tracking and data relay satellites

NASA Technical Reports Server (NTRS)

Oza, D. H.; Bolvin, D. T.; Lorah, J. M.; Lee, T.; Doll, C. E.

1995-01-01

The National Aeronautics and Space Administration (NASA) has developed the Tracking and Data Relay Satellite (TDRS) System (TDRSS) for tracking and communications support of low Earth-orbiting satellites. TDRSS has the operational capability of providing 85% coverage for TDRSS-user spacecraft. TDRSS currently consists of five geosynchronous spacecraft and the White Sands Complex (WSC) at White Sands, New Mexico. The Bilateration Ranging Transponder System (BRTS) provides range and Doppler measurements for each TDRS. The ground-based BRTS transponders are tracked as if they were TDRSS-user spacecraft. Since the positions of the BRTS transponders are known, their radiometric tracking measurements can be used to provide a well-determined ephemeris for the TDRS spacecraft. For high-accuracy orbit determination of a TDRSS user, such as the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft, high-accuracy TDRS orbits are required. This paper reports on successive refinements in improved techniques and procedures leading to more accurate TDRS orbit determination strategies using the Goddard Trajectory Determination System (GTDS). These strategies range from the standard operational solution using only the BRTS tracking measurements to a sophisticated iterative process involving several successive simultaneous solutions for multiple TDRSs and a TDRSS-user spacecraft. Results are presented for GTDS-generated TDRS ephemerides produced in simultaneous solutions with the TOPEX/Poseidon spacecraft. Strategies with different user spacecraft, as well as schemes for recovering accurate TDRS orbits following a TDRS maneuver, are also presented. In addition, a comprehensive assessment and evaluation of alternative strategies for TDRS orbit determination, excluding BRTS tracking measurements, are presented.

A strategy for developing a launch vehicle system for orbit insertion: Methodological aspects

NASA Astrophysics Data System (ADS)

Klyushnikov, V. Yu.; Kuznetsov, I. I.; Osadchenko, A. S.

2014-12-01

The article addresses methodological aspects of a development strategy to design a launch vehicle system for orbit insertion. The development and implementation of the strategy are broadly outlined. An analysis is provided of the criterial base and input data needed to define the main requirements for the launch vehicle system. Approaches are suggested for solving individual problems in working out the launch vehicle system development strategy.

Elastic issues and vibration reduction in a tethered deorbiting mission

NASA Astrophysics Data System (ADS)

Sabatini, Marco; Gasbarri, Paolo; Palmerini, Giovanni B.

2016-05-01

Recently proposed mission concepts involving harpoons or nets to capture and de-orbit debris represent an interesting application of the tethered systems, where the orbiting bodies are connected by a flexible link. These systems present a complex behavior, as flexible characteristics combine with orbital dynamics. The focus of the paper is on the dynamic behavior of the tethered system in the final phase of the de-orbiting mission, when a powerful apogee motor is used to change the debris orbit. The thrust action introduces significant issues, as elastic waves propagate along the tether, and the relevant oscillations couple with the orbital dynamics. Input shaping techniques are proposed to limit or cancel these oscillations. However, the performance of these techniques drops when non-ideal scenarios are considered. In particular, an initially slack tether is a serious issue that must be solved if acceptably low oscillations of the tether are to be obtained. Three strategies are proposed and discussed in this paper to remove the slack condition: a natural drift of the chaser by means of a single impulse, a controlled maneuver for precisely adjusting the relative distance between chaser spacecraft and debris, and a retrieval mechanism for changing the tether length.

Benign orbital apex tumors treated with multisession gamma knife radiosurgery.

PubMed

Goh, Alice Siew Ching; Kim, Yoon-Duck; Woo, Kyung In; Lee, Jung-Il

2013-03-01

The orbital apex is an important anatomic landmark that hosts numerous critical neurovascular structures. Tumor resection performed at this complex region poses a therapeutic challenge to orbital surgeons and often is associated with significant visual morbidity. This article reports the efficacy and safety of multisession gamma knife radiosurgery (GKRS) in benign, well-circumscribed tumors located at the orbital apex. Retrospective interventional case series. Five patients with visual disturbances resulting from a benign, well-circumscribed orbital apex tumor (3 cases of cavernous hemangioma and 2 cases of schwannoma). Each patient treated with GKRS with a total radiation dose of 20 Gy in 4 sessions (5 Gy in each session with an isodose line of 50%) delivered to the tumor margin. Best-corrected visual acuity, visual field changes, orbital imaging, tumor growth control, and side effects of radiation. All patients demonstrated improvement in visual acuity, pupillary responses, color vision, and visual field. Tumor shrinkage was observed in all patients and remained stable until the last follow-up. No adverse events were noted during or after the radiosurgery. None of the patients experienced any radiation-related ocular morbidity. From this experience, multisession GKRS seems to be an effective management strategy to treat solitary, benign, well-circumscribed orbital apex tumors. Copyright © 2013 American Academy of Ophthalmology. Published by Elsevier Inc. All rights reserved.

Low Earth Orbit Rendezvous Strategy for Lunar Missions

NASA Technical Reports Server (NTRS)

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

2006-01-01

On January 14, 2004 President George W. Bush announced a new Vision for Space Exploration calling for NASA to return humans to the moon. In 2005 NASA decided to use a Low Earth Orbit (LEO) rendezvous strategy for the lunar missions. A Discrete Event Simulation (DES) based model of this strategy was constructed. Results of the model were then used for subsequent analysis to explore the ramifications of the LEO rendezvous strategy.

Low-Latency Telerobotics from Mars Orbit: The Case for Synergy Between Science and Human Exploration

NASA Technical Reports Server (NTRS)

Valinia, A.; Garvin, J. B.; Vondrak, R.; Thronson, H.; Lester, D.; Schmidt, G.; Fong, T.; Wilcox, B.; Sellers, P.; White, N.

2012-01-01

Initial, science-directed human exploration of Mars will benefit from capabilities in which human explorers remain in orbit to control telerobotic systems on the surface (Figure 1). Low-latency, high-bandwidth telerobotics (LLT) from Mars orbit offers opportunities for what the terrestrial robotics community considers to be high-quality telepresence. Such telepresence would provide high quality sensory perception and situation awareness, and even capabilities for dexterous manipulation as required for adaptive, informed selection of scientific samples [1]. Astronauts on orbit in close communication proximity to a surface exploration site (in order to minimize communication latency) represent a capability that would extend human cognition to Mars (and potentially for other bodies such as asteroids, Venus, the Moon, etc.) without the challenges, expense, and risk of putting those humans on hazardous surfaces or within deep gravity wells. Such a strategy may be consistent with goals for a human space flight program that, are currently being developed within NASA.

Short rendezvous missions for advanced Russian human spacecraft

NASA Astrophysics Data System (ADS)

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.

Radiation exposure and performance of multiple burn LEO-GEO orbit transfer trajectories

NASA Technical Reports Server (NTRS)

Gorland, S. H.

1985-01-01

Many potential strategies exist for the transfer of spacecraft from low Earth orbit (LEO) to geosynchronous (GEO) orbit. One strategy has generally been utilized, that being a single impulsive burn at perigee and a GEO insertion burn at apogee. Multiple burn strategies were discussed for orbit transfer vehicles (OTVs) but the transfer times and radiation exposure, particularly for potentially manned missions, were used as arguments against those options. Quantitative results concerning the trip time and radiation encountered by multiple burn orbit transfer missions in order to establish the feasibility of manned missions, the vulnerability of electronics, and the shielding requirements are presented. The performance of these multiple burn missions is quantified in terms of the payload and propellant variances from the minimum energy mission transfer. The missions analyzed varied from one to eight perigee burns and ranged from a high thrust, 1 g acceleration, cryogenic hydrogen-oxygen chemical prpulsion system to a continuous burn, 0.001 g acceleration, hydrogen fueled resistojet propulsion system with a trip time of 60 days.

A guidance and control assessment of three vertical landing options for RLV

NASA Technical Reports Server (NTRS)

Gallaher, M.; Coughlin, D.; Krupp, D

1995-01-01

The National Aeronautics and Space Administration is considering a vertical lander as a candidate concept for a single-stage-to-orbit reusable launch vehicle (RLV). Three strategies for guiding and controlling the inversion of a reentering RLV from a nose-first attitude to a vertical landing attitude are suggested. Each option is simulated from a common reentry state to touchdown, using a common guidance algorithm and different controllers. Results demonstrate the characteristics that typify and distinguish each concept and help to identify peculiar problems, level of guidance and control sophistication required, feasibility concerns, and areas in which stringent subsystem requirements will be imposed by guidance and control.

[Orbital decompression in Grave's ophtalmopathy].

PubMed

Longueville, E

2010-01-01

Graves disease orbitopathy is a complex progressive inflammatory disease. Medical treatment remains in all cases the proposed treatment of choice. Surgical treatment by bone decompression can be considered as an emergency mainly in cases of optic neuropathy or ocular hypertension not being controlled medically or in post-traumatic exophthalmos stage. Emergency bone decompression eliminates compression or stretching of the optic nerve allowing visual recovery. The uncontrolled ocular hypertension will benefit from decompression. The normalization of intraocular pressure may be obtained by this surgery or if needed by the use of postoperative antiglaucoma drops or even filtration surgery. In all operated cases, the IOP was normalized with an average decrease of 7.71 mmHg and a cessation of eye drops in 3/7 cases. Regarding sequelae, our therapeutic strategy involves consecutively surgery of the orbit, extraocular muscles and eyelids. The orbital expansion gives excellent results on the cosmetic level and facilitates the implementation of subsequent actions.

Lattice Commissioning Stretgy Simulation for the B Factory

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

Lee, M.; Whittum, D.; Yan, Y.

2011-08-26

To prepare for the PEP-II turn on, we have studied one commissioning strategy with simulated lattice errors. Features such as difference and absolute orbit analysis and correction are discussed. To prepare for the commissioning of the PEP-II injection line and high energy ring (HER), we have developed a system for on-line orbit analysis by merging two existing codes: LEGO and RESOLVE. With the LEGO-RESOLVE system, we can study the problem of finding quadrupole alignment and beam position (BPM) offset errors with simulated data. We have increased the speed and versatility of the orbit analysis process by using a command filemore » written in a script language designed specifically for RESOLVE. In addition, we have interfaced the LEGO-RESOLVE system to the control system of the B-Factory. In this paper, we describe online analysis features of the LEGO-RESOLVE system and present examples of practical applications.« less

High-Payoff Space Transportation Design Approach with a Technology Integration Strategy

NASA Technical Reports Server (NTRS)

McCleskey, C. M.; Rhodes, R. E.; Chen, T.; Robinson, J.

2011-01-01

A general architectural design sequence is described to create a highly efficient, operable, and supportable design that achieves an affordable, repeatable, and sustainable transportation function. The paper covers the following aspects of this approach in more detail: (1) vehicle architectural concept considerations (including important strategies for greater reusability); (2) vehicle element propulsion system packaging considerations; (3) vehicle element functional definition; (4) external ground servicing and access considerations; and, (5) simplified guidance, navigation, flight control and avionics communications considerations. Additionally, a technology integration strategy is forwarded that includes: (a) ground and flight test prior to production commitments; (b) parallel stage propellant storage, such as concentric-nested tanks; (c) high thrust, LOX-rich, LOX-cooled first stage earth-to-orbit main engine; (d) non-toxic, day-of-launch-loaded propellants for upper stages and in-space propulsion; (e) electric propulsion and aero stage control.

Capture orbits around asteroids by hitting zero-velocity curves

NASA Astrophysics Data System (ADS)

Wang, Wei; Yang, Hongwei; Zhang, Wei; Ma, Guangfu

2017-12-01

The problem of capturing a spacecraft from a heliocentric orbit into a high parking orbit around binary asteroids is investigated in the current study. To reduce the braking Δ V, a new capture strategy takes advantage of the three-body gravity of the binary asteroid to lower the inertial energy before applying the Δ V. The framework of the circular restricted three-body problem (CR3BP) is employed for the binary asteroid system. The proposed capture strategy is based on the mechanism by which inertial energy can be decreased sharply near zero-velocity curves (ZVCs). The strategy has two steps, namely, hitting the target ZVC and raising the periapsis by a small Δ V at the apoapsis. By hitting the target ZVC, the positive inertial energy decreases and becomes negative. Using a small Δ V, the spacecraft inserts into a bounded orbit around the asteroid. In addition, a rotating mass dipole model is employed for elongated asteroids, which leads to dynamics similar to that of the CR3BP. With this approach, the proposed capture strategy can be applied to elongated asteroids. Numerical simulations validate that the proposed capture strategy is applicable for the binary asteroid 90 Antiope and the elongated asteroid 216 Kleopatra.

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

NASA Technical Reports Server (NTRS)

Barbee, Brent William; Carpenter, J. Russell; Heatwole, Scott; Markley, F. Landis; Moreau, Michael; Naasz, Bo J.; VanEepoel, John

2010-01-01

The feasibility and benefits of various spacecraft servicing concepts are currently being assessed, and all require that the servicer spacecraft perform rendezvous, proximity, and capture operations with the target spacecraft to be serviced. Many high-value spacecraft, which would be logical targets for servicing from an economic point of view, are located in geosynchronous orbit, a regime in which autonomous rendezvous and capture operations are not commonplace. Furthermore, existing GEO spacecraft were not designed to be serviced. Most do not have cooperative relative navigation sensors or docking features, and some servicing applications, such as de-orbiting of a non-functional spacecraft, entail rendezvous and capture with a spacecraft that may be non-functional or un-controlled. Several of these challenges have been explored via the design of a notional mission in which a nonfunctional satellite in geosynchronous orbit is captured by a servicer spacecraft and boosted into super-synchronous orbit for safe disposal. A strategy for autonomous rendezvous, proximity operations, and capture is developed, and the Orbit Determination Toolbox (ODTBX) is used to perform a relative navigation simulation to assess the feasibility of performing the rendezvous using a combination of angles-only and range measurements. Additionally, a method for designing efficient orbital rendezvous sequences for multiple target spacecraft is utilized to examine the capabilities of a servicer spacecraft to service multiple targets during the course of a single mission.

Project MEDSAT: The design of a remote sensing platform for malaria research and control

NASA Astrophysics Data System (ADS)

1991-04-01

Project MEDSAT was proposed with the specific goal of designing a satellite to remotely sense pertinent information useful in establishing strategies to control malaria. The 340 kg MEDSAT satellite is to be inserted into circular earth orbit aboard the Pegasus Air-Launched Space Booster at an inclination of 21 degrees and an altitude of 473 km. It is equipped with a synthetic aperture radar and a visible thermal/infrared sensor to remotely sense conditions at the target area of Chiapas, Mexico. The orbit is designed so that MEDSAT will pass over the target site twice each day. The data from each scan will be downlinked to Hawaii for processing, resulting in maps indicating areas of high malaria risk. These will be distributed to health officials at the target site. A relatively inexpensive launch by Pegasus and a design using mainly proven, off-the-shelf technology permit a low mission cost, while innovations in the satellite controls and the scientific instruments allow a fairly complex mission.

Project MEDSAT: The design of a remote sensing platform for malaria research and control

NASA Technical Reports Server (NTRS)

1991-01-01

Project MEDSAT was proposed with the specific goal of designing a satellite to remotely sense pertinent information useful in establishing strategies to control malaria. The 340 kg MEDSAT satellite is to be inserted into circular earth orbit aboard the Pegasus Air-Launched Space Booster at an inclination of 21 degrees and an altitude of 473 km. It is equipped with a synthetic aperture radar and a visible thermal/infrared sensor to remotely sense conditions at the target area of Chiapas, Mexico. The orbit is designed so that MEDSAT will pass over the target site twice each day. The data from each scan will be downlinked to Hawaii for processing, resulting in maps indicating areas of high malaria risk. These will be distributed to health officials at the target site. A relatively inexpensive launch by Pegasus and a design using mainly proven, off-the-shelf technology permit a low mission cost, while innovations in the satellite controls and the scientific instruments allow a fairly complex mission.

On-Orbit Propulsion and Methods of Momentum Management for the International Space Station

NASA Technical Reports Server (NTRS)

Russell, Samuel P.; Spencer, Victor; Metrocavage, Kevin; Swanson, Robert A.; Krajchovich, Mark; Beisner, Matthew; Kamath, Ulhas P.

2010-01-01

Since the first documented design of a space station in 1929, it has been a dream of many to sustain a permanent presence in space. Russia and the US spent several decades competing for a sustained human presence in low Earth orbit. In the 1980 s, Russia and the US began to openly collaborate to achieve this goal. This collaboration lead to the current design of the ISS. Continuous improvement of procedures for controlling the ISS have lead to more efficient propellant management over the years. Improved efficiency combined with the steady use of cargo vehicles has kept ISS propellant levels well above their defined thresholds in all categories. The continuing evolution of propellant and momentum management operational strategies demonstrates the capability and flexibility of the ISS propulsion system. The hard work and cooperation of the international partners and the evolving operational strategies have made the ISS safe and successful. The ISS s proven success is the foundation for the future of international cooperation for sustaining life in space.

Neural network-based sliding mode control for atmospheric-actuated spacecraft formation using switching strategy

NASA Astrophysics Data System (ADS)

Sun, Ran; Wang, Jihe; Zhang, Dexin; Shao, Xiaowei

2018-02-01

This paper presents an adaptive neural networks-based control method for spacecraft formation with coupled translational and rotational dynamics using only aerodynamic forces. It is assumed that each spacecraft is equipped with several large flat plates. A coupled orbit-attitude dynamic model is considered based on the specific configuration of atmospheric-based actuators. For this model, a neural network-based adaptive sliding mode controller is implemented, accounting for system uncertainties and external perturbations. To avoid invalidation of the neural networks destroying stability of the system, a switching control strategy is proposed which combines an adaptive neural networks controller dominating in its active region and an adaptive sliding mode controller outside the neural active region. An optimal process is developed to determine the control commands for the plates system. The stability of the closed-loop system is proved by a Lyapunov-based method. Comparative results through numerical simulations illustrate the effectiveness of executing attitude control while maintaining the relative motion, and higher control accuracy can be achieved by using the proposed neural-based switching control scheme than using only adaptive sliding mode controller.

Operationally optimal maneuver strategy for spacecraft injected into sub-geosynchronous transfer orbit

NASA Astrophysics Data System (ADS)

Kiran, B. S.; Singh, Satyendra; Negi, Kuldeep

The GSAT-12 spacecraft is providing Communication services from the INSAT/GSAT system in the Indian region. The spacecraft carries 12 extended C-band transponders. GSAT-12 was launched by ISRO’s PSLV from Sriharikota, into a sub-geosynchronous Transfer Orbit (sub-GTO) of 284 x 21000 km with inclination 18 deg. This Mission successfully accomplished combined optimization of launch vehicle and satellite capabilities to maximize operational life of the s/c. This paper describes mission analysis carried out for GSAT-12 comprising launch window, orbital events study and orbit raising maneuver strategies considering various Mission operational constraints. GSAT-12 is equipped with two earth sensors (ES), three gyroscopes and digital sun sensor. The launch window was generated considering mission requirement of minimum 45 minutes of ES data for calibration of gyros with Roll-sun-pointing orientation in T.O. Since the T.O. period was a rather short 6.1 hr, required pitch biases were worked out to meet the gyro-calibration requirement. A 440 N Liquid Apogee Motor (LAM) is used for orbit raising. The objective of the maneuver strategy is to achieve desired drift orbit satisfying mission constraints and minimizing propellant expenditure. In case of sub-GTO, the optimal strategy is to first perform an in-plane maneuver at perigee to raise the apogee to synchronous level and then perform combined maneuvers at the synchronous apogee to achieve desired drift orbit. The perigee burn opportunities were examined considering ground station visibility requirement for monitoring the burn. Two maneuver strategies were proposed: an optimal five-burn strategy with two perigee burns centered around perigee#5 and perigee#8 with partial ground station visibility and three apogee burns with dual station visibility, a near-optimal five-burn strategy with two off-perigee burns at perigee#5 and perigee#8 with single ground station visibility and three apogee burns with dual station visibility. The range vector profiles were studied in the s/c frame during LAM burn phases and accurate polarization predictions were provided to supporting ground stations. The near optimal strategy was selected for implementation in order to ensure full visibility during each LAM burn. Contingency maneuver plans were generated in preparation for specified Propulsion system related contingencies. Maneuver plans were generated considering 3-sigma dispersions in T.O. GSAT-12 is positioned at 83 deg East longitude. The estimated operational life is about 11 years which was realized through operationally optimal maneuver strategy selected from the detailed mission analysis.

Navigating highly elliptical earth orbiters with simultaneous VLBI from orthogonal baseline pairs

NASA Technical Reports Server (NTRS)

Frauenholz, Raymond B.

1986-01-01

Navigation strategies for determining highly elliptical orbits with VLBI are described. The predicted performance of wideband VLBI and Delta VLBI measurements obtained by orthogonal baseline pairs are compared for a 16-hr equatorial orbit. It is observed that the one-sigma apogee position accuracy improves two orders of magnitude to the meter level when Delta VLBI measurements are added to coherent Doppler and range, and the simpler VLBI strategy provides nearly the same orbit accuracy. The effects of differential measurement noise and acquisition geometry on orbit accuracy are investigated. The data reveal that quasar position uncertainty limits the accuracy of wideband Delta VLBI measurements, and that polar motion and baseline uncertainties and offsets between station clocks affect the wideband VLBI data. It is noted that differential one-way range (DOR) has performance nearly equal to that of the more complex Delta DOR and is recommended for use on spacecraft in high elliptical orbits.

The end-of-life disposal of satellites in libration-point orbits using solar radiation pressure

NASA Astrophysics Data System (ADS)

Soldini, Stefania; Colombo, Camilla; Walker, Scott

2016-04-01

This paper proposes an end-of-life propellant-free disposal strategy for libration-point orbits which uses solar radiation pressure to restrict the evolution of the spacecraft motion. The spacecraft is initially disposed into the unstable manifold leaving the libration-point orbit, before a reflective sun-pointing surface is deployed to enhance the effect of solar radiation pressure. Therefore, the consequent increase in energy prevents the spacecraft's return to Earth. Three European Space Agency missions are selected as test case scenarios: Herschel, SOHO and Gaia. Guidelines for the end-of-life disposal of future libration-point orbit missions are proposed and a preliminary study on the effect of the Earth's orbital eccentricity on the disposal strategy is shown for the Gaia mission.

Horizon: A Proposal for Large Aperture, Active Optics in Geosynchronous Orbit

NASA Technical Reports Server (NTRS)

Chesters, Dennis; Jenstrom, Del

2000-01-01

In 1999, NASA's New Millennium Program called for proposals to validate new technology in high-earth orbit for the Earth Observing-3 (NMP EO3) mission to fly in 2003. In response, we proposed to test a large aperture, active optics telescope in geosynchronous orbit. This would flight-qualify new technologies for both Earth and Space science: 1) a future instrument with LANDSAT image resolution and radiometric quality watching continuously from geosynchronous station, and 2) the Next Generation Space Telescope (NGST) for deep space imaging. Six enabling technologies were to be flight-qualified: 1) a 3-meter, lightweight segmented primary mirror, 2) mirror actuators and mechanisms, 3) a deformable mirror, 4) coarse phasing techniques, 5) phase retrieval for wavefront control during stellar viewing, and 6) phase diversity for wavefront control during Earth viewing. Three enhancing technologies were to be flight- validated: 1) mirror deployment and latching mechanisms, 2) an advanced microcontroller, and 3) GPS at GEO. In particular, two wavefront sensing algorithms, phase retrieval by JPL and phase diversity by ERIM International, were to sense optical system alignment and focus errors, and to correct them using high-precision mirror mechanisms. Active corrections based on Earth scenes are challenging because phase diversity images must be collected from extended, dynamically changing scenes. In addition, an Earth-facing telescope in GEO orbit is subject to a powerful diurnal thermal and radiometric cycle not experienced by deep-space astronomy. The Horizon proposal was a bare-bones design for a lightweight large-aperture, active optical system that is a practical blend of science requirements, emerging technologies, budget constraints, launch vehicle considerations, orbital mechanics, optical hardware, phase-determination algorithms, communication strategy, computational burdens, and first-rate cooperation among earth and space scientists, engineers and managers. This manuscript presents excerpts from the Horizon proposal's sections that describe the Earth science requirements, the structural -thermal-optical design, the wavefront sensing and control, and the on-orbit validation.

A study of an orbital radar mapping mission to Venus. Volume 3: Parametric studies and subsystem comparisons

NASA Technical Reports Server (NTRS)

1973-01-01

Parametric studies and subsystem comparisons for the orbital radar mapping mission to planet Venus are presented. Launch vehicle requirements and primary orbiter propulsion system requirements are evaluated. The systems parametric analysis indicated that orbit size and orientation interrelated with almost all of the principal spacecraft systems and influenced significantly the definition of orbit insertion propulsion requirements, weight in orbit capability, radar system design, and mapping strategy.

Shuttle on-orbit rendezvous targeting: Circular orbits

NASA Technical Reports Server (NTRS)

Bentley, E. L.

1972-01-01

The strategy and logic used in a space shuttle on-orbit rendezvous targeting program are described. The program generates ascent targeting conditions for boost to insertion into an intermediate parking orbit, and generates on-orbit targeting and timeline bases for each maneuver to effect rendezvous with a space station. Time of launch is determined so as to eliminate any plane change, and all work was performed for a near-circular space station orbit.

Contamination control concepts for space station customer servicing

NASA Technical Reports Server (NTRS)

Maruya, K. A.; Ryan, L. E.; Rosales, L. A.; Medler, E. H.

1986-01-01

The customer servicing operations envisioned for the space station, which include instrument repair, orbital replacement unit (ORU) changeout, and fluid replenishment for free-flying and attached payloads, are expected to create requirements for a unique contamination control subsystem for the customer servicing facility (CSF). Both the core space station and the CSF users present unique requirements/sensitivities, not all of which are currently defined with common criteria. Preliminary results from an assessment of the effects of the CSF-induced contamination environment are reported. Strategies for a comprehensive contamination control approach and a description of specific hardware devices and their applicability are discussed.

Maintaining Aura's Orbit Requirements While Performing Orbit Maintenance Maneuvers Containing an Orbit Normal Delta-V Component

NASA Technical Reports Server (NTRS)

Johnson, Megan R.; Petersen, Jeremy D.

2014-01-01

The Earth Observing System (EOS) Afternoon Constellation consists of five member missions (GCOM-W1, Aqua, CALIPSO, CloudSat, and Aura), each of which maintain a frozen, sun-synchronous orbit with a 16-day repeating ground track that follows the Worldwide Reference System-2 (WRS-2). Under nominal science operations for Aura, the propulsion system is oriented such that the resultant thrust vector is aligned 13.493 degrees away from the velocity vector along the yaw axis. When performing orbit maintenance maneuvers, the spacecraft performs a yaw slew to align the thrust vector in the appropriate direction. A new Drag Make Up (DMU) maneuver operations scheme has been implemented for Aura alleviating the need for the 13.493 degree yaw slew. The focus of this investigation is to assess the impact that no-slew DMU maneuver operations will have on Aura's Mean Local Time (MLT) which drives the required along track separation between Aura and the constellation members, as well as Aura's frozen orbit properties, eccentricity and argument of perigee. Seven maneuver strategies were analyzed to determine the best operational approach. A mirror pole strategy, with maneuvers alternating at the North and South poles, was implemented operationally to minimize impact to the MLT. Additional analysis determined that the mirror pole strategy could be further modified to include frozen orbit maneuvers and thus maintain both MLT and the frozen orbit properties under noslew operations.

Best Mitigation Paths To Effectively Reduce Earth's Orbital Debris

NASA Technical Reports Server (NTRS)

Wiegman, Bruce M.

2009-01-01

This slide presentation reviews some ways to reduce the problem posed by debris in orbit around the Earth. It reviews the orbital debris environment, the near-term needs to minimize the Kessler syndrome, also known as collisional cascading, a survey of active orbital debris mitigation strategies, the best paths to actively remove orbital debris, and technologies that are required for active debris mitigation.

Precise orbit determination and rapid orbit recovery supported by time synchronization

NASA Astrophysics Data System (ADS)

Guo, Rui; Zhou, JianHua; Hu, XiaoGong; Liu, Li; Tang, Bo; Li, XiaoJie; Wu, Shan

2015-06-01

In order to maintain optimal signal coverage, GNSS satellites have to experience orbital maneuvers. For China's COMPASS system, precise orbit determination (POD) as well as rapid orbit recovery after maneuvers contribute to the overall Positioning, Navigation and Timing (PNT) service performance in terms of accuracy and availability. However, strong statistical correlations between clock offsets and the radial component of a satellite's positions require long data arcs for POD to converge. We propose here a new strategy which relies on time synchronization between ground tracking stations and in-orbit satellites. By fixing satellite clock offsets measured by the satellite station two-way synchronization (SSTS) systems and receiver clock offsets, POD and orbital recovery performance can be improved significantly. Using the Satellite Laser Ranging (SLR) as orbital accuracy evaluation, we find the 4-hr recovered orbit achieves about 0.71 m residual root mean square (RMS) error of fit SLR data, the recovery time is improved from 24-hr to 4-hr compared with the conventional POD without time synchronization support. In addition, SLR evaluation shows that for 1-hr prediction, about 1.47 m accuracy is achieved with the new proposed POD strategy.

Bottom-up synthesis of WS2 nanosheets with synchronous surface modification for imaging guided tumor regression.

PubMed

Wang, Shige; Zhao, Jiulong; Yang, Hailun; Wu, Chenyao; Hu, Fei; Chang, Haizhou; Li, Guixiang; Ma, Dan; Zou, Duowu; Huang, Mingxian

2017-08-01

Two-dimensional transition metal dichalcogenides (TMDs) have been receiving great attention as NIR photothermal transducing agent in tumor photothermal therapy. Keeping in mind the low efficiency of the conventional top-down exfoliated 2D TMDs and the complexity of their surface modifications, we herein proposed a bottom-up strategy for the one-pot hydrothermal and controlled synthesis of surface polyvinyl pyrrolidone (PVP) modified WS 2 nanosheets. The material design was based on the chelating-coordinating effect between the lone pair electrons of oxygen of PVP carbonyl group and the unoccupied orbital (5d orbitals) of tungsten. The WS 2 nanosheets with synchronous surface PVP grafting showed an excellent photothermal conversion performance, while the surface anchored PVP guaranteed its colloidal stability. Moreover, the strong X-ray attenuation ability and near-infrared (NIR) absorbance of WS 2 -PVP 360kDa enabled the sensitive in vitro and in vivo computed tomography and photoacoustic imaging. The WS 2 -PVP 360kDa nanosheets were biocompatible and exhibited promising in vitro and in vivo anti-cancer efficacy. Findings in this report may greatly promote the design of colloidal stable and biocompatible 2D TMDs and their future clinical translations. A bottom-up strategy for the one-pot and controlled synthesis of surface polyvinyl pyrrolidone (PVP) modified WS 2 nanosheets was proposed for the first time. By hydrothermally treating the mixture solution of tetrathiotungstate and PVP, Owing to the chelating-coordinating effect between the lone pair electrons of oxygen of PVP carbonyl group and the unoccupied orbital (5d orbitals) of tungsten, PVP was synchronously graphed on WS 2 -PVP nanosheets surface. The formed WS 2 -PVP nanosheets were colloidal stable, biocompatible, and exhibited promising computed tomography, photoacoustic imaging and tumor photothermal therapy efficacy both in vitro and in vivo. Copyright © 2017 Acta Materialia Inc. Published by Elsevier Ltd. All rights reserved.

Mars Reconnaissance Orbiter Navigation Strategy for Mars Science Laboratory Entry, Descent and Landing Telecommunication Relay Support

NASA Technical Reports Server (NTRS)

Williams, Jessica L.; Menon, Premkumar R.; Demcak, Stuart W.

2012-01-01

The Mars Reconnaissance Orbiter (MRO) is an orbiting asset that performs remote sensing observations in order to characterize the surface, subsurface and atmosphere of Mars. To support upcoming NASA Mars Exploration Program Office objectives, MRO will be used as a relay communication link for the Mars Science Laboratory (MSL) mission during the MSL Entry, Descent and Landing sequence. To do so, MRO Navigation must synchronize the MRO Primary Science Orbit (PSO) with a set of target conditions requested by the MSL Navigation Team; this may be accomplished via propulsive maneuvers. This paper describes the MRO Navigation strategy for and operational performance of MSL EDL relay telecommunication support.

Cislunar Near Rectilinear Halo Orbit for Human Space Exploration

NASA Technical Reports Server (NTRS)

Whitley, Ryan; Martinez, Roland; Condon, Gerald; Williams, Jacob; Lee, David; Davis, Diane; Barton, Gregg; Bhatt, Sagar; Jang, Jiann-Woei; Clark, Fred;

Spacecraft stability and control using new techniques for periodic and time-delayed systems

NASA Astrophysics Data System (ADS)

NAzari, Morad

This dissertation addresses various problems in spacecraft stability and control using specialized theoretical and numerical techniques for time-periodic and time-delayed systems. First, the effects of energy dissipation are considered in the dual-spin spacecraft, where the damper masses in the platform (?) and the rotor (?) cause energy loss in the system. Floquet theory is employed to obtain stability charts for different relative spin rates of the subsystem [special characters omitted] with respect to the subsystem [special characters omitted]. Further, the stability and bifurcation of delayed feedback spin stabilization of a rigid spacecraft is investigated. The spin is stabilized about the principal axis of the intermediate moment of inertia using a simple delayed feedback control law. In particular, linear stability is analyzed via the exponential-polynomial characteristic equations and then the method of multiple scales is used to obtain the normal form of the Hopf bifurcation. Next, the dynamics of a rigid spacecraft with nonlinear delayed multi-actuator feedback control are studied, where a nonlinear feedback controller using an inverse dynamics approach is sought for the controlled system to have the desired linear delayed closed-loop dynamics (CLD). Later, three linear state feedback control strategies based on Chebyshev spectral collocation and the Lyapunov Floquet transformation (LFT) are explored for regulation control of linear periodic time delayed systems. First , a delayed feedback control law with discrete delay is implemented and the stability of the closed-loop response is investigated in the parameter space of available control gains using infinite-dimensional Floquet theory. Second, the delay differential equation (DDE) is discretized into a large set of ordinary differential equations (ODEs) using the Chebyshev spectral continuous time approximation (CSCTA) and delayed feedback with distributed delay is applied. The third strategy involves use of both CSCTA and the reduced Lyapunov Floquet transformation (RLFT) in order to design a non-delayed feedback control law. The delayed Mathieu equation is used as an illustrative example in which the closed-loop response and control effort are compared for all three control strategies. Finally, three example applications of control of time-periodic astrodynamic systems, i.e. formation flying control for an elliptic Keplerian chief orbit, body-fixed hovering control over a tumbling asteroid, and stationkeeping in Earth-Moon L1 halo orbits, are shown using versions of the control strategies introduced above. These applications employ a mixture of feedforward and non-delayed periodic-gain state feedback for tracking control of natural and non-natural motions in these systems. A major conclusion is that control effort is minimized by employing periodic-gain (rather than constant-gain) feedback control in such systems.

Orbit-associated tumors: navigation and control of resection using intraoperative computed tomography.

PubMed

Terpolilli, Nicole A; Rachinger, Walter; Kunz, Mathias; Thon, Niklas; Flatz, Wilhelm H; Tonn, Jörg-Christian; Schichor, Christian

2016-05-01

OBJECT Treatment of skull base lesions is complex and usually requires a multidisciplinary approach. In meningioma, which is the most common tumor entity in this region, resection is considered to be the most important therapeutic step to avoid tumor recurrence. However, resection of skull base lesions with orbital or optic nerve involvement poses a challenge due to their anatomical structure and their proximity to eloquent areas. Therefore the main goal of surgery should be to achieve the maximum extent of resection while preserving neurological function. In the postoperative course, medical and radiotherapeutic strategies may then be successfully used to treat possible tumor residues. Methods to safely improve the extent of resection in skull base lesions therefore are desirable. The current study reports the authors' experience with the use of intraoperative CT (iCT) combined with neuronavigation with regard to feasibility and possible benefits of the method. METHODS Those patients with tumorous lesions in relationship to the orbit, sphenoid wing, or cavernous sinus who were surgically treated between October 2008 and December 2013 using iCT-based neuronavigation and in whom an intraoperative scan was obtained for control of resection were included. In all cases a second iCT scan was performed under sterile conditions after completion of navigation-guided microsurgical tumor resection. The surgical strategy was adapted accordingly; if necessary, resection was continued. RESULTS Twenty-three patients (19 with WHO Grade I meningioma and 4 with other lesions) were included. The most common clinical symptoms were loss of visual acuity and exophthalmus. Intraoperative control of resection by iCT was successfully obtained in all cases. Intraoperative imaging changed the surgical approach in more than half (52.2%) of these patients, either because iCT demonstrated unexpected residual tumor masses or because the second scan revealed additional tumor tissue that was not detected in the first scan due to overlay by osseous tumor parts; in these cases resection was continued. In the remaining cases resection was concluded as planned because iCT verified the surgeon's microscopic estimation of tumor resection status. Postoperative visual outcome was favorable in more than 80% of patients. CONCLUSIONS Intraoperative CT allows control of resection in case of uncertainty and can help to improve the extent of maximal safe resection, especially in case of osseous tumor parts and masses within the orbit.

Using Mean Orbit Period in Mars Reconnaissance Orbiter Maneuver Design

NASA Technical Reports Server (NTRS)

Chung, Min-Kun J.; Menon, Premkumar R.; Wagner, Sean V.; Williams, Jessica L.

2014-01-01

Mars Reconnaissance Orbiter (MRO) has provided communication relays for a number of Mars spacecraft. In 2016 MRO is expected to support a relay for NASA's Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) spacecraft. In addition, support may be needed by another mission, ESA's ExoMars EDL Demonstrator Module's (EDM), only 21 days after the InSight coverage. The close proximity of these two events presents a unique challenge to a conventional orbit synchronization maneuver where one deterministic maneuver is executed prior to each relay. Since the two events are close together and the difference in required phasing between InSight and EDM may be up to half an orbit (yielding a large execution error), the downtrack timing error can increase rapidly at the EDM encounter. Thus, a new maneuver strategy that does not require a deterministic maneuver in-between the two events (with only a small statistical cleanup) is proposed in the paper. This proposed strategy rests heavily on the stability of the mean orbital period. The ability to search and set the specified mean period is fundamental in the proposed maneuver design as well as in understanding the scope of the problem. The proposed strategy is explained and its result is used to understand and solve the problem in the flight operations environment.

Planning and Control for Microassembly of Structures Composed of Stress-Engineered MEMS Microrobots

PubMed Central

Donald, Bruce R.; Levey, Christopher G.; Paprotny, Igor; Rus, Daniela

2013-01-01

We present control strategies that implement planar microassembly using groups of stress-engineered MEMS microrobots (MicroStressBots) controlled through a single global control signal. The global control signal couples the motion of the devices, causing the system to be highly underactuated. In order for the robots to assemble into arbitrary planar shapes despite the high degree of underactuation, it is desirable that each robot be independently maneuverable (independently controllable). To achieve independent control, we fabricated robots that behave (move) differently from one another in response to the same global control signal. We harnessed this differentiation to develop assembly control strategies, where the assembly goal is a desired geometric shape that can be obtained by connecting the chassis of individual robots. We derived and experimentally tested assembly plans that command some of the robots to make progress toward the goal, while other robots are constrained to remain in small circular trajectories (closed-loop orbits) until it is their turn to move into the goal shape. Our control strategies were tested on systems of fabricated MicroStressBots. The robots are 240–280 μm × 60 μm × 7–20 μm in size and move simultaneously within a single operating environment. We demonstrated the feasibility of our control scheme by accurately assembling five different types of planar microstructures. PMID:23580796

Advanced Navigation Strategies for an Asteroid Sample Return Mission

NASA Technical Reports Server (NTRS)

Bauman, J.; Getzandanner, K.; Williams, B.; Williams, K.

2011-01-01

The proximity operations phases of a sample return mission to an asteroid have been analyzed using advanced navigation techniques derived from experience gained in planetary exploration. These techniques rely on tracking types such as Earth-based radio metric Doppler and ranging, spacecraft-based ranging, and optical navigation using images of landmarks on the asteroid surface. Navigation strategies for the orbital phases leading up to sample collection, the touch down for collecting the sample, and the post sample collection phase at the asteroid are included. Options for successfully executing the phases are studied using covariance analysis and Monte Carlo simulations of an example mission to the near Earth asteroid 4660 Nereus. Two landing options were studied including trajectories with either one or two bums from orbit to the surface. Additionally, a comparison of post-sample collection strategies is presented. These strategies include remaining in orbit about the asteroid or standing-off a given distance until departure to Earth.

Analytical solutions to optimal underactuated spacecraft formation reconfiguration

NASA Astrophysics Data System (ADS)

Huang, Xu; Yan, Ye; Zhou, Yang

2015-11-01

Underactuated systems can generally be defined as systems with fewer number of control inputs than that of the degrees of freedom to be controlled. In this paper, analytical solutions to optimal underactuated spacecraft formation reconfiguration without either the radial or the in-track control are derived. By using a linear dynamical model of underactuated spacecraft formation in circular orbits, controllability analysis is conducted for either underactuated case. Indirect optimization methods based on the minimum principle are then introduced to generate analytical solutions to optimal open-loop underactuated reconfiguration problems. Both fixed and free final conditions constraints are considered for either underactuated case and comparisons between these two final conditions indicate that the optimal control strategies with free final conditions require less control efforts than those with the fixed ones. Meanwhile, closed-loop adaptive sliding mode controllers for both underactuated cases are designed to guarantee optimal trajectory tracking in the presence of unmatched external perturbations, linearization errors, and system uncertainties. The adaptation laws are designed via a Lyapunov-based method to ensure the overall stability of the closed-loop system. The explicit expressions of the terminal convergent regions of each system states have also been obtained. Numerical simulations demonstrate the validity and feasibility of the proposed open-loop and closed-loop control schemes for optimal underactuated spacecraft formation reconfiguration in circular orbits.

Space station orbit maintenance

NASA Technical Reports Server (NTRS)

Kaplan, D. I.; Jones, R. M.

1983-01-01

The orbit maintenance problem is examined for two low-earth-orbiting space station concepts - the large, manned Space Operations Center (SOC) and the smaller, unmanned Science and Applications Space Platform (SASP). Atmospheric drag forces are calculated, and circular orbit altitudes are selected to assure a 90 day decay period in the event of catastrophic propulsion system failure. Several thrusting strategies for orbit maintenance are discussed. Various chemical and electric propulsion systems for orbit maintenance are compared on the basis of propellant resupply requirements, power requirements, Shuttle launch costs, and technology readiness.

Integrated propulsion for near-Earth space missions. Volume 2: Technical

NASA Technical Reports Server (NTRS)

Dailey, C. L.; Meissinger, H. F.; Lovberg, R. H.; Zafran, S.

1981-01-01

The calculation approach is described for parametric analysis of candidate electric propulsion systems employed in LEO to GEO missions. Occultation relations, atmospheric density effects, and natural radiation effects are presented. A solar cell cover glass tradeoff is performed to determine optimum glass thickness. Solar array and spacecraft pointing strategies are described for low altitude flight and for optimum array illumination during ascent. Mass ratio tradeoffs versus transfer time provide direction for thruster technology improvements. Integrated electric propulsion analysis is performed for orbit boosting, inclination change, attitude control, stationkeeping, repositioning, and disposal functions as well as power sharing with payload on orbit. Comparison with chemical auxiliary propulsion is made to quantify the advantages of integrated propulsion in terms of weight savings and concomittant launch cost savings.

System identification for modeling for control of flexible structures

NASA Technical Reports Server (NTRS)

Mettler, Edward; Milman, Mark

1986-01-01

The major components of a design and operational flight strategy for flexible structure control systems are presented. In this strategy an initial distributed parameter control design is developed and implemented from available ground test data and on-orbit identification using sophisticated modeling and synthesis techniques. The reliability of this high performance controller is directly linked to the accuracy of the parameters on which the design is based. Because uncertainties inevitably grow without system monitoring, maintaining the control system requires an active on-line system identification function to supply parameter updates and covariance information. Control laws can then be modified to improve performance when the error envelopes are decreased. In terms of system safety and stability the covariance information is of equal importance as the parameter values themselves. If the on-line system ID function detects an increase in parameter error covariances, then corresponding adjustments must be made in the control laws to increase robustness. If the error covariances exceed some threshold, an autonomous calibration sequence could be initiated to restore the error enveloped to an acceptable level.

Management of the orbital environment

NASA Technical Reports Server (NTRS)

Loftus, Joseph P., Jr.; Kessler, Donald J.; Anz-Meador, Phillip D.

1991-01-01

Data regarding orbital debris are presented to shed light on the requirements of environmental management in space, and strategies are given for active intervention and operational strategies. Debris are generated by inadvertent explosions of upper stages, intentional military explosions, and collisional breakups. Design and operation practices are set forth for minimizing debris generation and removing useless debris from orbit in the low-earth and geosynchronous orbits. Self-disposal options include propulsive maneuvers, drag-augmentation devices, and tether systems, and the drag devices are described as simple and passive. Active retrieval and disposition are considered, and the difficulty is examined of removing small debris. Active intervention techniques are required since pollution prevention is more effective than remediation for the problems of both earth and space.

Accessibility, stabilizability, and feedback control of continuous orbital transfer.

PubMed

Gurfil, Pini

2004-05-01

This paper investigates the problem of low-thrust orbital transfer using orbital element feedback from a control-theoretic standpoint, concepts of controllability, feedback stabilizability, and their interaction. The Gauss variational equations (GVEs) are used to model the state-space dynamics. First, the notion of accessibility, a weaker form of controllability, is presented. It is then shown that the GVEs are globally accessible. Based on the accessibility result, a nonlinear feedback controller is derived that asymptotically steers a vehicle from an initial elliptic Keplerian orbit to any given elliptic Keplerian orbit. The performance of the new controller is illustrated by simulating an orbital transfer between two geosynchronous Earth orbits. It is shown that the low-thrust controller requires less fuel than an impulsive maneuver for the same transfer time. Closed-form, analytic expressions for the new orbital transfer controller are given. Finally, it is proved, based on a topological nonlinear stabilizability test, that there does not exist a continuous closed-loop controller that can transfer a spacecraft to a parabolic escape trajectory.

Control of nonlinear flexible space structures

NASA Astrophysics Data System (ADS)

Shi, Jianjun

With the advances made in computer technology and efficiency of numerical algorithms over last decade, the MPC strategies have become quite popular among control community. However, application of MPC or GPC to flexible space structure control has not been explored adequately in the literature. The work presented in this thesis primarily focuses on application of GPC to control of nonlinear flexible space structures. This thesis is particularly devoted to the development of various approximate dynamic models, design and assessment of candidate controllers, and extensive numerical simulations for a realistic multibody flexible spacecraft, namely, Jupiter Icy Moons Orbiter (JIMO)---a Prometheus class of spacecraft proposed by NASA for deep space exploratory missions. A stable GPC algorithm is developed for Multi-Input-Multi-Output (MIMO) systems. An end-point weighting (penalty) is used in the GPC cost function to guarantee the nominal stability of the closed-loop system. A method is given to compute the desired end-point state from the desired output trajectory. The methodologies based on Fake Algebraic Riccati Equation (FARE) and constrained nonlinear optimization, are developed for synthesis of state weighting matrix. This makes this formulation more practical. A stable reconfigurable GPC architecture is presented and its effectiveness is demonstrated on both aircraft as well as spacecraft model. A representative in-orbit maneuver is used for assessing the performance of various control strategies using various design models. Different approximate dynamic models used for analysis include linear single body flexible structure, nonlinear single body flexible structure, and nonlinear multibody flexible structure. The control laws evaluated include traditional GPC, feedback linearization-based GPC (FLGPC), reconfigurable GPC, and nonlinear dissipative control. These various control schemes are evaluated for robust stability and robust performance in the presence of parametric uncertainties and input disturbances. Finally, the conclusions are made with regard to the efficacy of these controllers and potential directions for future research.

Orbit determination and prediction of GEO satellite of BeiDou during repositioning maneuver

NASA Astrophysics Data System (ADS)

Cao, Fen; Yang, XuHai; Li, ZhiGang; Sun, BaoQi; Kong, Yao; Chen, Liang; Feng, Chugang

2014-11-01

In order to establish a continuous GEO satellite orbit during repositioning maneuvers, a suitable maneuver force model has been established associated with an optimal orbit determination method and strategy. A continuous increasing acceleration is established by constructing a constant force that is equivalent to the pulse force, with the mass of the satellite decreasing throughout maneuver. This acceleration can be added to other accelerations, such as solar radiation, to obtain the continuous acceleration of the satellite. The orbit determination method and strategy are illuminated, with subsequent assessment of the orbit being determined and predicted accordingly. The orbit of the GEO satellite during repositioning maneuver can be determined and predicted by using C-Band pseudo-range observations of the BeiDou GEO satellite with COSPAR ID 2010-001A in 2011 and 2012. The results indicate that observations before maneuver do affect orbit determination and prediction, and should therefore be selected appropriately. A more precise orbit and prediction can be obtained compared to common short arc methods when observations starting 1 day prior the maneuver and 2 h after the maneuver are adopted in POD (Precise Orbit Determination). The achieved URE (User Range Error) under non-consideration of satellite clock errors is better than 2 m within the first 2 h after maneuver, and less than 3 m for further 2 h of orbit prediction.

Galileo Jupiter approach orbit determination

NASA Technical Reports Server (NTRS)

Miller, J. K.; Nicholson, F. T.

1984-01-01

Orbit determination characteristics of the Jupiter approach phase of the Galileo mission are described. Predicted orbit determination performance is given for the various mission events that occur during Jupiter approach. These mission events include delivery of an atmospheric entry probe, acquisition of probe science data by the Galileo orbiter for relay to earth, delivery of an orbiter to a close encounter of the Galilean satellite Io, and insertion of the orbiter into orbit about Jupiter. The orbit determination strategy and resulting accuracies are discussed for the data types which include Doppler, range, optical imaging of Io, and a new Very Long Baseline Interferometry (VLBI) data type called Differential One-Way Range (DOR).

Versatile low-Reynolds-number swimmer with three-dimensional maneuverability.

PubMed

Jalali, Mir Abbas; Alam, Mohammad-Reza; Mousavi, SeyyedHossein

2014-11-01

We design and simulate the motion of a swimmer, the Quadroar, with three-dimensional translation and reorientation capabilities in low-Reynolds-number conditions. The Quadroar is composed of an I-shaped frame whose body link is a simple linear actuator and four disks that can rotate about the axes of flange links. The time symmetry is broken by a combination of disk rotations and the one-dimensional expansion or contraction of the body link. The Quadroar propels on forward and transverse straight lines and performs full three-dimensional reorientation maneuvers, which enable it to swim along arbitrary trajectories. We find continuous operation modes that propel the swimmer on planar and three-dimensional periodic and quasiperiodic orbits. Precessing quasiperiodic orbits consist of slow lingering phases with cardioid or multiloop turns followed by directional propulsive phases. Quasiperiodic orbits allow the swimmer to access large parts of its neighboring space without using complex control strategies. We also discuss the feasibility of fabricating a nanoscale Quadroar by photoactive molecular rotors.

PSA: A program to streamline orbit determination for launch support operations

NASA Technical Reports Server (NTRS)

Legerton, V. N.; Mottinger, N. A.

1988-01-01

An interactive, menu driven computer program was written to streamline the orbit determination process during the critical launch support phase of a mission. Residing on a virtual memory minicomputer, this program retains the quantities in-core needed to obtain a least squares estimate of the spacecraft trajectory with interactive displays to assist in rapid radio metric data evaluation. Menu-driven displays allow real time filter and data strategy development. Graphical and tabular displays can be sent to a laser printer for analysis without exiting the program. Products generated by this program feed back to the main orbit determination program in order to further refine the estimate of the trajectory. The final estimate provides a spacecraft ephemeris which is transmitted to the mission control center and used for antenna pointing and frequency predict generation by the Deep Space Network. The development and implementation process of this program differs from that used for most other navigation software by allowing the users to check important operating features during development and have changes made as needed.

A strategy for Earth science from space in the 1980s. Part 1: Solid earth and oceans

NASA Technical Reports Server (NTRS)

1982-01-01

The report develops a ten-year science strategy for investigating the solid earth and dynamics of world oceans from Earth orbit. The strategy begins from the premise that earth studies have proceeded to the point where further advances in understanding Earth processes must be based on a global perspective and that the U.S. is technically ready to begin a global study approach from Earth orbit. The major areas of study and their fundamental problems are identified. The strategy defines the primary science objectives to be addressed and the essential measurements and precision to achieve them.

Guidance and Control strategies for aerospace vehicles

NASA Technical Reports Server (NTRS)

Hibey, J. L.; Naidu, D. S.; Charalambous, C. D.

1989-01-01

A neighboring optimal guidance scheme was devised for a nonlinear dynamic system with stochastic inputs and perfect measurements as applicable to fuel optimal control of an aeroassisted orbital transfer vehicle. For the deterministic nonlinear dynamic system describing the atmospheric maneuver, a nominal trajectory was determined. Then, a neighboring, optimal guidance scheme was obtained for open loop and closed loop control configurations. Taking modelling uncertainties into account, a linear, stochastic, neighboring optimal guidance scheme was devised. Finally, the optimal trajectory was approximated as the sum of the deterministic nominal trajectory and the stochastic neighboring optimal solution. Numerical results are presented for a typical vehicle. A fuel-optimal control problem in aeroassisted noncoplanar orbital transfer is also addressed. The equations of motion for the atmospheric maneuver are nonlinear and the optimal (nominal) trajectory and control are obtained. In order to follow the nominal trajectory under actual conditions, a neighboring optimum guidance scheme is designed using linear quadratic regulator theory for onboard real-time implementation. One of the state variables is used as the independent variable in reference to the time. The weighting matrices in the performance index are chosen by a combination of a heuristic method and an optimal modal approach. The necessary feedback control law is obtained in order to minimize the deviations from the nominal conditions.

Combined Orbital Fractures: Surgical Strategy of Sequential Repair

PubMed Central

Hur, Su Won; Kim, Sung Eun; Chung, Kyu Jin; Lee, Jun Ho; Kim, Tae Gon

2015-01-01

Background Reconstruction of combined orbital floor and medial wall fractures with a comminuted inferomedial strut (IMS) is challenging and requires careful practice. We present our surgical strategy and postoperative outcomes. Methods We divided 74 patients who underwent the reconstruction of the orbital floor and medial wall concomitantly into a comminuted IMS group (41 patients) and non-comminuted IMS group (33 patients). In the comminuted IMS group, we first reconstructed the floor stably and then the medial wall by using separate implant pieces. In the non-comminuted IMS group, we reconstructed the floor and the medial wall with a single large implant. Results In the follow-up of 6 to 65 months, most patients with diplopia improved in the first-week except one, who eventually improved at 1 year. All patients with an EOM limitation improved during the first month of follow-up. Enophthalmos (displacement, 2 mm) was observed in two patients. The orbit volume measured on the CT scans was statistically significantly restored in both groups. No complications related to the surgery were observed. Conclusions We recommend the reconstruction of orbit walls in the comminuted IMS group by using the following surgical strategy: usage of multiple pieces of rigid implants instead of one large implant, sequential repair first of the floor and then of the medial wall, and a focus on the reconstruction of key areas. Our strategy of step-by-step reconstruction has the benefits of easy repair, less surgical trauma, and minimal stress to the surgeon. PMID:26217562

The performance of differential VLBI delay during interplanetary cruise

NASA Technical Reports Server (NTRS)

Moultrie, B.; Wolff, P. J.; Taylor, T. H.

1984-01-01

Project Voyager radio metric data are used to evaluate the orbit determination abilities of several data strategies during spacecraft interplanetary cruise. Benchmark performance is established with an operational data strategy of conventional coherent doppler, coherent range, and explicitly differenced range data from two intercontinental baselines to ameliorate the low declination singularity of the doppler data. Employing a Voyager operations trajectory as a reference, the performance of the operational data strategy is compared to the performances of data strategies using differential VLBI delay data (spacecraft delay minus quasar delay) in combinations with the aforementioned conventional data types. The comparison of strategy performances indicates that high accuracy cruise orbit determination can be achieved with a data strategy employing differential VLBI delay data, where the quantity of coherent radio metric data has been greatly reduced.

Control of the Spin Angular Momentum and Orbital Angular Momentum of a Reflected Wave by Multifunctional Graphene Metasurfaces.

PubMed

Zhang, Chen; Deng, Li; Zhu, Jianfeng; Hong, Weijun; Wang, Ling; Yang, Wenjie; Li, Shufang

2018-06-21

Three kinds of multifunctional graphene metasurfaces based on Pancharatnam⁻Berry (PB) phase cells are proposed and numerically demonstrated to control a reflected wave’s spin angular momentum (SAM) and orbital angular momentum (OAM) in the terahertz (THz) regime. Each proposed metasurface structure is composed of an array of graphene strips with different deviation angles and a back-grounded quartz substrate. In order to further help readers have a deeper insight into the graphene-based metasurfaces, a detailed design strategy is also provided. With the aid of the designed graphene elements, the proposed metasurfaces can achieve the full 360° range of phase coverage and provide manipulation of SAM and OAM of a circularly polarized (CP) wave at will. More importantly, simultaneous control of these two momentums can also be realized, and in order to demonstrate this function, a THz spin-controlled OAM beam generator with diverse topological charges is created, which can provide one more degree of freedom to improve the channel capability without increasing the bandwidth compared to a linearly polarized (LP) OAM beam. Numerical results verify the proposed graphene metasurfaces, which pave the way for generating spin OAM vortex waves for THz communication systems.

Autonomous docking system for space structures and satellites

NASA Astrophysics Data System (ADS)

Prasad, Guru; Tajudeen, Eddie; Spenser, James

2005-05-01

Aximetric proposes Distributed Command and Control (C2) architecture for autonomous on-orbit assembly in space with our unique vision and sensor driven docking mechanism. Aximetric is currently working on ip based distributed control strategies, docking/mating plate, alignment and latching mechanism, umbilical structure/cord designs, and hardware/software in a closed loop architecture for smart autonomous demonstration utilizing proven developments in sensor and docking technology. These technologies can be effectively applied to many transferring/conveying and on-orbit servicing applications to include the capturing and coupling of space bound vehicles and components. The autonomous system will be a "smart" system that will incorporate a vision system used for identifying, tracking, locating and mating the transferring device to the receiving device. A robustly designed coupler for the transfer of the fuel will be integrated. Advanced sealing technology will be utilized for isolation and purging of resulting cavities from the mating process and/or from the incorporation of other electrical and data acquisition devices used as part of the overall smart system.

Management of complex orbital fractures.

PubMed

Bhatti, N; Kanzaria, A; Huxham-Owen, N; Bridle, C; Holmes, S

2016-09-01

The treatment of orbital injuries has evolved considerably over the last two decades. We describe strategies involved in the emergency management of orbital injuries, the use of imaging, preformed and customised materials for reconstruction, and endoscopic techniques. Copyright © 2016 The British Association of Oral and Maxillofacial Surgeons. Published by Elsevier Ltd. All rights reserved.

Systematic Identification of Preferred Orbits for Magnetospheric Missions. 1; Single Satellites

NASA Technical Reports Server (NTRS)

Stern, David P.

2000-01-01

This is a systematic attempt to identify and assess near-equatorial, high-eccentricity orbits best suited for studying the Earth's magnetosphere, in particular its most dynamic part, the plasma sheet of the magnetotail. The study was motivated by the design needs of a multi-spacecraft "constellation" mission, stressing low cost, minimal active control and economic launch strategies, and both quantitative and qualitative aspects were investigated. On one hand, by collecting hourly samples throughout the year, accurate estimates were obtained of the coverage of different regions, and of the frequency and duration of long eclipses. On the other hand, an intuitive understanding was developed of the factors which determine the merits of the mission, including long-range factors due to perturbations by the Moon, Sun and the Earth's equatorial bulge.

Precise orbit determination of Multi-GNSS constellation including GPS GLONASS BDS and GALIEO

NASA Astrophysics Data System (ADS)

Dai, Xiaolei

2014-05-01

In addition to the existing American global positioning system (GPS) and the Russian global navigation satellite system (GLONASS), the new generation of GNSS is emerging and developing, such as the Chinese BeiDou satellite navigation system (BDS) and the European GALILEO system. Multi-constellation is expected to contribute to more accurate and reliable positioning and navigation service. However, the application of multi-constellation challenges the traditional precise orbit determination (POD) strategy that was designed usually for single constellation. In this contribution, we exploit a more rigorous multi-constellation POD strategy for the ongoing IGS multi-GNSS experiment (MGEX) where the common parameters are identical for each system, and the frequency- and system-specified parameters are employed to account for the inter-frequency and inter-system biases. Since the authorized BDS attitude model is not yet released, different BDS attitude model are implemented and their impact on orbit accuracy are studied. The proposed POD strategy was implemented in the PANDA (Position and Navigation Data Analyst) software and can process observations from GPS, GLONASS, BDS and GALILEO together. The strategy is evaluated with the multi-constellation observations from about 90 MGEX stations and BDS observations from the BeiDou experimental tracking network (BETN) of Wuhan University (WHU). Of all the MGEX stations, 28 stations record BDS observation, and about 80 stations record GALILEO observations. All these data were processed together in our software, resulting in the multi-constellation POD solutions. We assessed the orbit accuracy for GPS and GLONASS by comparing our solutions with the IGS final orbit, and for BDS and GALILEO by overlapping our daily orbit solution. The stability of inter-frequency bias of GLONASS and inter-system biases w.r.t. GPS for GLONASS, BDS and GALILEO were investigated. At last, we carried out precise point positioning (PPP) using the multi-constellation POD orbit and clock products, and analyzed the contribution of these POD products to PPP. Keywords: Multi-GNSS, Precise Orbit Determination, Inter-frequency bias, Inter-system bias, Precise Point Positioning

Fronto-orbital feminization technique. A surgical strategy using fronto-orbital burring with or without eggshell technique to optimize the risk/benefit ratio.

PubMed

Villepelet, A; Jafari, A; Baujat, B

2018-05-04

The demand for facial feminization is increasing in transsexual patients. Masculine foreheads present extensive supraorbital bossing with a more acute glabellar angle, whereas female foreheads show softer features. The aim of this article is to describe our surgical technique for fronto-orbital feminization. The mask-lift technique is an upper face-lift. It provides rejuvenation by correcting collapsed features, and fronto-orbital feminization through burring of orbital rims and lateral canthopexies. Depending on the size of the frontal sinus and the thickness of its anterior wall, frontal remodeling is achieved using simple burring or by means of the eggshell technique. Orbital remodeling comprises a superolateral orbital opening, a reduction of ridges and a trough at the lateral orbital rim to support the lateral canthopexy. Frontal, corrugator and procerus myectomies, plus minimal scalp excision, complete the surgery. Our technique results in significant, natural-looking feminization. No complications were observed in our series of patients. The eggshell technique is an alternative to bone flap on over-pneumatized sinus. Fronto-orbital feminization fits into a wider surgical strategy. It can be associated to rhinoplasty, genioplasty, mandibular angle remodeling, face lift and laryngoplasty. Achieving facial feminization in 2 or 3 stages improves psychological and physiological tolerance. Copyright © 2018 Elsevier Masson SAS. All rights reserved.

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

NASA Astrophysics Data System (ADS)

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 activation procedure for a de-orbit strategy and the level of integration it should have with the host satellite in order to be activated before a total failure. At Cranfield Space Research Centre two different solutions have already been developed as de-orbit sail payloads for microsatellites (Icarus-1 on TechDemoSat-1 and Icarus-3 on Carbonite-1 currently on-orbit, DOM for future ESA ESEO mission). This study will provide a useful input to improve and refine the current de-orbit concepts for future satellite missions.

Nondynamic Tracking Using The Global Positioning System

NASA Technical Reports Server (NTRS)

Yunck, T. P.; Wu, Sien-Chong

1988-01-01

Report describes technique for using Global Positioning System (GPS) to determine position of low Earth orbiter without need for dynamic models. Differential observing strategy requires GPS receiver on user vehicle and network of six ground receivers. Computationally efficient technique delivers decimeter accuracy on orbits down to lowest altitudes. New technique nondynamic long-arc strategy having potential for accuracy of best dynamic techniques while retaining much of computational simplicity of geometric techniques.

Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance

NASA Technical Reports Server (NTRS)

Long, Stacia M.; You, Tung-Han; Halsell, C. Allen; Bhat, Ramachand S.; Demcak, Stuart W.; Graat, Eric J.; Higa, Earl S.; Highsmith, Dolan E.; Mottinger, Neil A.; Jah, Moriba K.

2007-01-01

The Mars Reconnaissance Orbiter reached Mars on March 10, 2006 and performed a Mars orbit insertion maneuver of 1 km/s to enter into a large elliptical orbit. Three weeks later, aerobraking operations began and lasted about five months. Aerobraking utilized the atmospheric drag to reduce the large elliptical orbit into a smaller, near circular orbit. At the time of MRO aerobraking, there were three other operational spacecraft orbiting Mars and the navigation team had to minimize the possibility of a collision. This paper describes the daily operations of the MRO navigation team during this time as well as the collision avoidance strategy development and implementation.

Mars Observer Lecture: Mars Orbit Insertion

NASA Technical Reports Server (NTRS)

Dodd, Suzanne R. (Personal Name)

1993-01-01

The Mars Observer mission spacecraft was primarily designed for exploring Mars and the Martian environment. The Mars Observer was launched on September 25, 1992. The spacecraft was lost in the vicinity of Mars on August 21, 1993 when the spacecraft began its maneuvering sequence for Martian orbital insertion. This videotape shows a lecture by Suzanne R. Dodd, the Mission Planning Team Chief for the Mars Observer Project. Ms Dodd begins with a brief overview of the mission and the timeline from the launch to orbital insertion. Ms Dodd then reviews slides showing the trajectory of the spacecraft on its trip to Mars. Slides of the spacecraft being constructed are also shown. She then discusses the Mars orbit insertion and the events that will occur to move the spacecraft from the capture orbit into a mapping orbit. During the trip to Mars, scientists at JPL had devised a new strategy, called Power In that would allow for an earlier insertion into the mapping orbit. The talk summarizes this strategy, showing on a slide the planned transition orbits. There are shots of the Martian moon, Phobos, taken from the Viking spacecraft, as Ms Dodd explains that the trajectory will allow the orbiter to make new observations of that moon. She also explains the required steps to prepare for mapping after the spacecraft has achieved the mapping orbit around Mars. The lecture ends with a picture of Mars from the Observer on its approach to the planet.

An analytic algorithm for global coverage of the revisiting orbit and its application to the CFOSAT satellite

NASA Astrophysics Data System (ADS)

Xu, Ming; Huang, Li

2014-08-01

This paper addresses a new analytic algorithm for global coverage of the revisiting orbit and its application to the mission revisiting the Earth within long periods of time, such as Chinese-French Oceanic Satellite (abbr., CFOSAT). In the first, it is presented that the traditional design methodology of the revisiting orbit for some imaging satellites only on the single (ascending or descending) pass, and the repeating orbit is employed to perform the global coverage within short periods of time. However, the selection of the repeating orbit is essentially to yield the suboptimum from the rare measure of rational numbers of passes per day, which will lose lots of available revisiting orbits. Thus, an innovative design scheme is proposed to check both rational and irrational passes per day to acquire the relationship between the coverage percentage and the altitude. To improve the traditional imaging only on the single pass, the proposed algorithm is mapping every pass into its ascending and descending nodes on the specified latitude circle, and then is accumulating the projected width on the circle by the field of view of the satellite. The ergodic geometry of coverage percentage produced from the algorithm is affecting the final scheme, such as the optimal one owning the largest percentage, and the balance one possessing the less gradient in its vicinity, and is guiding to heuristic design for the station-keeping control strategies. The application of CFOSAT validates the feasibility of the algorithm.

A comparative study between control strategies for a solar sailcraft in an Earth-Mars transfer

NASA Astrophysics Data System (ADS)

Mainenti-Lopes, I.; Souza, L. C. Gadelha; De Sousa, Fabiano. L.

2016-10-01

The goal of this work was a comparative study of solar sail trajectory optimization using different control strategies. Solar sailcraft is propulsion system with great interest in space engineering, since it uses solar radiation to propulsion. So there is no need for propellant to be used, thus it can remains active throughout the entire transfer maneuver. This type of propulsion system opens the possibility to reduce the cost of exploration missions in the solar system. In its simplest configuration, a Flat Solar Sail (FSS) consists of a large and thin structure generally composed by a film fixed to flexible rods. The performance of these vehicles depends largely on the sails attitude relative to the Sun. Using a FSS as propulsion, an Earth-Mars transfer optimization problem was tackled by the algorithms GEOreal1 and GEOreal2 (Generalized Extremal Optimization with real codification). Those algorithms are Evolutionary Algorithms (AE) based on the theory of Self-Organized Criticality. They were used to optimize the FSS attitude angle so it could reach Mars orbit in minimum time. It was considered that the FSS could perform up to ten attitude maneuvers during orbital transfer. Moreover, the time between maneuvers can be different. So, the algorithms had to optimize an objective function with 20 design variables. The results obtained in this work were compared with previously results that considered constant values of time between maneuvers.

On-orbit flight control algorithm description

NASA Technical Reports Server (NTRS)

1975-01-01

Algorithms are presented for rotational and translational control of the space shuttle orbiter in the orbital mission phases, which are external tank separation, orbit insertion, on-orbit and de-orbit. The program provides a versatile control system structure while maintaining uniform communications with other programs, sensors, and control effectors by using an executive routine/functional subroutine format. Software functional requirements are described using block diagrams where feasible, and input--output tables, and the software implementation of each function is presented in equations and structured flow charts. Included are a glossary of all symbols used to define the requirements, and an appendix of supportive material.

The ENVISAT Atmospheric Chemistry mission (GOMOS, MIPAS and SCIAMACHY) - Instrument status and mission evolution

NASA Astrophysics Data System (ADS)

Niro, F.

2009-04-01

The ENVISAT ESA's satellite was launched on a polar orbit on March 2002. It carries on-board three atmospheric chemistry instruments: GOMOS, MIPAS and SCIAMACHY [1]. At the present time, although the mission expected lifetime of 5 years has been already exceeded, all the payload modules are in good to excellent status. The only limiting factor is the available fuel that is used for orbit control manoeuvre. Recently a new strategy was proposed [2] that will allow to save fuel and to extend the mission up to 2013. Following this strategy, the altitude of the orbit will be lowered by 17 km starting from end of 2010 and the inclination will be allowed to drift. The new orbit scenario will result in a new repeating cycle with a variation of the Mean Local Solar Time (MLST). This will have an impact on both the in-flight operations, on the science data and on the mission. The simulations carried out for the atmospheric chemistry instruments show that the new orbit strategy will neither have a significant impact in the instrument operations nor on the quality of the science data. Therefore we expect that the atmospheric mission will continue nominally until the end of the platform life time, providing to the scientist a unique dataset of the most important geophysical parameters (e.g., trace gases, clouds, and aerosol) spanning a time interval of about 11 years. The aim of this paper is to review the overall ENVISAT atmospheric mission status for the past, present and future. The evolution of the instrument performances since launch will be analyzed with focus on the life-limited items monitoring. The tuning of the instrument in-flight operations decided to cope with instrument degradation or scientific needs will be described. The lessons learned on how to operate and monitor the instruments will be highlighted. Finally the expected evolution of the instrument performances until the ENVISAT end-of-life will be discussed. [1] H. Nett, J. Frerick, T. Paulsen, and G. Levrini, "The atmospheric instruments and their applications: GOMOS, MIPAS and SCIAMACHY", ESA Bulletin (ISSN 0376-4265), No. 106, p. 77 - 87 (2001) [2] J. Frerick, B. Duesmann, and M. Canela, "2010 and beyond - The ENVISAT mission extension", Proc. ‘Envisat Symposium 2007', Montreux, Switzerland, 23-27 April 2007 (ESA SP-636, July 2007)

The ENVISAT Atmospheric Chemistry mission (GOMOS, MIPAS and SCIAMACHY) -Instrument status and mission evolution

NASA Astrophysics Data System (ADS)

Dehn, Angelika

The ENVISAT ESA's satellite was launched on a polar orbit on March 2002. It carries on-board three atmospheric chemistry instruments: GOMOS, MIPAS and SCIAMACHY [1]. At the present time, although the mission expected lifetime of 5 years has been already exceeded, all the payload modules are in good to excellent status. The only limiting factor is the available fuel that is used for orbit control manoeuvre. A new strategy was proposed [2] that will allow to save fuel and to extend the mission up to 2013. Following this strategy, the altitude of the orbit will be lowered by 17 km starting from end of 2010 and the inclination will be allowed to drift. The new orbit scenario will result in a new repeating cycle with a variation of the Mean Local Solar Time (MLST). This will have an impact on both the in-flight operations, on the science data and on the mission. The simulations carried out for the atmospheric chemistry instruments show that the new orbit strategy will neither have a significant impact in the instrument operations nor on the quality of the science data. Therefore we expect that the atmospheric mission will continue nominally until the end of the platform life time, providing to the scientist a unique dataset of the most important geophysical parameters (e.g., trace gases, clouds, and aerosol) spanning a time interval of about 11 years. The aim of this paper is to review the overall ENVISAT atmospheric mission status for the past, present and future. The evolution of the instrument performances since launch will be analyzed with focus on the life-limited items monitoring. The tuning of the instrument in-flight operations decided to cope with instrument degradation or scientific needs will be described. The lessons learned on how to operate and monitor the instruments will be highlighted. Finally the expected evolution of the instrument performances until the ENVISAT end-of-life will be discussed. [1] H. Nett, J. Frerick, T. Paulsen, and G. Levrini, "The atmospheric instruments and their applications: GOMOS, MIPAS and SCIAMACHY", ESA Bulletin (ISSN 0376-4265), No. 106, p. 77 -87 (2001) [2] J. Frerick, B. Duesmann, and M. Canela, "2010 and beyond -The ENVISAT mission extension", Proc. `Envisat Symposium 2007', Montreux, Switzerland, 23-27 April 2007 (ESA SP-636, July 2007)

A space exploration strategy that promotes international and commercial participation

NASA Astrophysics Data System (ADS)

Arney, Dale C.; Wilhite, Alan W.; Chai, Patrick R.; Jones, Christopher A.

2014-01-01

NASA has created a plan to implement the Flexible Path strategy, which utilizes a heavy lift launch vehicle to deliver crew and cargo to orbit. In this plan, NASA would develop much of the transportation architecture (launch vehicle, crew capsule, and in-space propulsion), leaving the other in-space elements open to commercial and international partnerships. This paper presents a space exploration strategy that reverses that philosophy, where commercial and international launch vehicles provide launch services. Utilizing a propellant depot to aggregate propellant on orbit, smaller launch vehicles are capable of delivering all of the mass necessary for space exploration. This strategy has benefits to the architecture in terms of cost, schedule, and reliability.

Plant growth strategies are remodeled by spaceflight.

PubMed

Paul, Anna-Lisa; Amalfitano, Claire E; Ferl, Robert J

2012-12-07

Arabidopsis plants were grown on the International Space Station within specialized hardware that combined a plant growth habitat with a camera system that can capture images at regular intervals of growth. The Imaging hardware delivers telemetric data from the ISS, specifically images received in real-time from experiments on orbit, providing science without sample return. Comparable Ground Controls were grown in a sister unit that is maintained in the Orbital Environment Simulator at Kennedy Space Center. One of many types of biological data that can be analyzed in this fashion is root morphology. Arabidopsis seeds were geminated on orbit on nutrient gel Petri plates in a configuration that encouraged growth along the surface of the gel. Photos were taken every six hours for the 15 days of the experiment. In the absence of gravity, but the presence of directional light, spaceflight roots remained strongly negatively phototropic and grew in the opposite direction of the shoot growth; however, cultivars WS and Col-0 displayed two distinct, marked differences in their growth patterns. First, cultivar WS skewed strongly to the right on orbit, while cultivar Col-0 grew with little deviation away from the light source. Second, the Spaceflight environment also impacted the rate of growth in Arabidopsis. The size of the Flight plants (as measured by primary root and hypocotyl length) was uniformly smaller than comparably aged Ground Control plants in both cultivars. Skewing and waving, thought to be gravity dependent phenomena, occur in spaceflight plants. In the presence of an orienting light source, phenotypic trends in skewing are gravity independent, and the general patterns of directional root growth typified by a given genotype in unit gravity are recapitulated on orbit, although overall growth patterns on orbit are less uniform. Skewing appears independent of axial orientation on the ISS - suggesting that other tropisms (such as for oxygen and temperature) do not influence skewing. An aspect of the spaceflight environment also retards the rate of early Arabidopsis growth.

Plant growth strategies are remodeled by spaceflight

PubMed Central

2012-01-01

Background Arabidopsis plants were grown on the International Space Station within specialized hardware that combined a plant growth habitat with a camera system that can capture images at regular intervals of growth. The Imaging hardware delivers telemetric data from the ISS, specifically images received in real-time from experiments on orbit, providing science without sample return. Comparable Ground Controls were grown in a sister unit that is maintained in the Orbital Environment Simulator at Kennedy Space Center. One of many types of biological data that can be analyzed in this fashion is root morphology. Arabidopsis seeds were geminated on orbit on nutrient gel Petri plates in a configuration that encouraged growth along the surface of the gel. Photos were taken every six hours for the 15 days of the experiment. Results In the absence of gravity, but the presence of directional light, spaceflight roots remained strongly negatively phototropic and grew in the opposite direction of the shoot growth; however, cultivars WS and Col-0 displayed two distinct, marked differences in their growth patterns. First, cultivar WS skewed strongly to the right on orbit, while cultivar Col-0 grew with little deviation away from the light source. Second, the Spaceflight environment also impacted the rate of growth in Arabidopsis. The size of the Flight plants (as measured by primary root and hypocotyl length) was uniformly smaller than comparably aged Ground Control plants in both cultivars. Conclusions Skewing and waving, thought to be gravity dependent phenomena, occur in spaceflight plants. In the presence of an orienting light source, phenotypic trends in skewing are gravity independent, and the general patterns of directional root growth typified by a given genotype in unit gravity are recapitulated on orbit, although overall growth patterns on orbit are less uniform. Skewing appears independent of axial orientation on the ISS – suggesting that other tropisms (such as for oxygen and temperature) do not influence skewing. An aspect of the spaceflight environment also retards the rate of early Arabidopsis growth. PMID:23217113

Space Operations Center orbit altitude selection strategy

NASA Technical Reports Server (NTRS)

Indrikis, J.; Myers, H. L.

1982-01-01

The strategy for the operational altitude selection has to respond to the Space Operation Center's (SOC) maintenance requirements and the logistics demands of the missions to be supported by the SOC. Three orbit strategies are developed: two are constant altitude, and one variable altitude. In order to minimize the effect of atmospheric uncertainty the dynamic altitude method is recommended. In this approach the SOC will operate at the optimum altitude for the prevailing atmospheric conditions and logistics model, provided that mission safety constraints are not violated. Over a typical solar activity cycle this method produces significant savings in the overall logistics cost.

A Collision Avoidance Strategy for a Potential Natural Satellite around the Asteroid Bennu for the OSIRIS-REx Mission

NASA Technical Reports Server (NTRS)

Mashiku, Alinda K.; Carpenter, J. Russell

2016-01-01

The cadence of proximity operations for the OSIRIS-REx mission may have an extra induced challenge given the potential of the detection of a natural satellite orbiting the asteroid Bennu. Current ground radar observations for object detection orbiting Bennu show no found objects within bounds of specific size and rotation rates. If a natural satellite is detected during approach, a different proximity operation cadence will need to be implemented as well as a collision avoidance strategy for mission success. A collision avoidance strategy will be analyzed using the Wald Sequential Probability Ratio Test.

A Collision Avoidance Strategy for a Potential Natural Satellite Around the Asteroid Bennu for the OSIRIS-REx Mission

NASA Technical Reports Server (NTRS)

Mashiku, Alinda; Carpenter, Russell

2016-01-01

The cadence of proximity operations for the OSIRIS-REx mission may have an extra induced challenge given the potential of the detection of a natural satellite orbiting the asteroid Bennu. Current ground radar observations for object detection orbiting Bennu show no found objects within bounds of specific size and rotation rates. If a natural satellite is detected during approach, a different proximity operation cadence will need to be implemented as well as a collision avoidance strategy for mission success. A collision avoidance strategy will be analyzed using the Wald Sequential Probability Ratio Test.

Report Of The HST Strategy Panel: A Strategy For Recovery

DTIC Science & Technology

1991-01-01

orbit change out: the Wide Field/Planetary Camera II (WFPC II), the Near-Infrared Camera and Multi- Object Spectrometer (NICMOS) and the Space ...are the Space Telescope Imaging Spectrograph (STB), the Near-Infrared Camera and Multi- Object Spectrom- eter (NICMOS), and the second Wide Field and...expected to fail to lock due to duplicity was 20%; on- orbit data indicates that 10% may be a better estimate, but the guide stars were preselected

The Role of GRAIL Orbit Determination in Preprocessing of Gravity Science Measurements

NASA Technical Reports Server (NTRS)

Kruizinga, Gerhard; Asmar, Sami; Fahnestock, Eugene; Harvey, Nate; Kahan, Daniel; Konopliv, Alex; Oudrhiri, Kamal; Paik, Meegyeong; Park, Ryan; Strekalov, Dmitry;

The Exploration of Mars Launch and Assembly Simulation

NASA Technical Reports Server (NTRS)

Cates, Grant; Stromgren, Chel; Mattfeld, Bryan; Cirillo, William; Goodliff, Kandyce

2016-01-01

Advancing human exploration of space beyond Low Earth Orbit, and ultimately to Mars, is of great interest to NASA, other organizations, and space exploration advocates. Various strategies for getting to Mars have been proposed. These include NASA's Design Reference Architecture 5.0, a near-term flyby of Mars advocated by the group Inspiration Mars, and potential options developed for NASA's Evolvable Mars Campaign. Regardless of which approach is used to get to Mars, they all share a need to visualize and analyze their proposed campaign and evaluate the feasibility of the launch and on-orbit assembly segment of the campaign. The launch and assembly segment starts with flight hardware manufacturing and ends with final departure of a Mars Transfer Vehicle (MTV), or set of MTVs, from an assembly orbit near Earth. This paper describes a discrete event simulation based strategic visualization and analysis tool that can be used to evaluate the launch campaign reliability of any proposed strategy for exploration beyond low Earth orbit. The input to the simulation can be any manifest of multiple launches and their associated transit operations between Earth and the exploration destinations, including Earth orbit, lunar orbit, asteroids, moons of Mars, and ultimately Mars. The simulation output includes expected launch dates and ascent outcomes i.e., success or failure. Running 1,000 replications of the simulation provides the capability to perform launch campaign reliability analysis to determine the probability that all launches occur in a timely manner to support departure opportunities and to deliver their payloads to the intended orbit. This allows for quantitative comparisons between alternative scenarios, as well as the capability to analyze options for improving launch campaign reliability. Results are presented for representative strategies.

Breaking the current density threshold in spin-orbit-torque magnetic random access memory

NASA Astrophysics Data System (ADS)

Zhang, Yin; Yuan, H. Y.; Wang, X. S.; Wang, X. R.

2018-04-01

Spin-orbit-torque magnetic random access memory (SOT-MRAM) is a promising technology for the next generation of data storage devices. The main bottleneck of this technology is the high reversal current density threshold. This outstanding problem is now solved by a new strategy in which the magnitude of the driven current density is fixed while the current direction varies with time. The theoretical limit of minimal reversal current density is only a fraction (the Gilbert damping coefficient) of the threshold current density of the conventional strategy. The Euler-Lagrange equation for the fastest magnetization reversal path and the optimal current pulse is derived for an arbitrary magnetic cell and arbitrary spin-orbit torque. The theoretical limit of minimal reversal current density and current density for a GHz switching rate of the new reversal strategy for CoFeB/Ta SOT-MRAMs are, respectively, of the order of 105 A/cm 2 and 106 A/cm 2 far below 107 A/cm 2 and 108 A/cm 2 in the conventional strategy. Furthermore, no external magnetic field is needed for a deterministic reversal in the new strategy.

Design and Stability of an On-Orbit Attitude Control System Using Reaction Control Thrusters

NASA Technical Reports Server (NTRS)

Hall, Robert A.; Hough, Steven; Orphee, Carolina; Clements, Keith

2016-01-01

NASA is providing preliminary design and requirements for the Space Launch System Exploration Upper Stage (EUS). The EUS will provide upper stage capability for vehicle ascent as well as on-orbit control capability. Requirements include performance of on-orbit burn to provide Orion vehicle with escape velocity. On-orbit attitude control is accommodated by a on-off Reaction Control System (RCS). Paper provides overview of approaches for design and stability of an attitude control system using a RCS.

Neurovestibular Considerations for Sub-Orbital Space Flight: A Framework for Future Investigation

PubMed Central

Karmali, Faisal; Shelhamer, Mark

2013-01-01

Commercial sub-orbital operators will soon offer the excitement of traveling to space to thousands of people. Based on previous experience in space flight and parabolic flight, sensorimotor disruptions in eye movements, postural stability, and motor coordination are likely in these travelers. Here we propose a framework for developing strategies to overcome these sensorimotor disruptions. We delineate how approaches should differ from those applied to orbital flight and between sub-orbital passengers and pilots based on differing frequency of flights and mission objectives. Sensorimotor adaptation is one strategy for overcoming disruptions; an important question is whether it occurs quickly enough to be of use during periods of reduced and enhanced gravity lasting less than five minutes. Data are presented showing that sensorimotor adaptation of the pitch vestibulo-ocular reflex during parabolic flight takes a few consecutive days of flying to overcome an initial disruption. We conclude with recommendations for operators and researchers to improve safety and comfort during sub-orbital operations. We recommend using parabolic flight as a tool for pre-adapting sub-orbital passengers, along with further research into the required quantity and timing of these pre-adaptation flights and the tasks conducted during these flights. Likewise, for sub-orbital pilots, we recommend emphasizing recency of experience. PMID:20555165

Practical method to identify orbital anomaly as spacecraft breakup in the geostationary region

NASA Astrophysics Data System (ADS)

Hanada, Toshiya; Uetsuhara, Masahiko; Nakaniwa, Yoshitaka

2012-07-01

Identifying a spacecraft breakup is an essential issue to define the current orbital debris environment. This paper proposes a practical method to identify an orbital anomaly, which appears as a significant discontinuity in the observation data, as a spacecraft breakup. The proposed method is applicable to orbital anomalies in the geostationary region. Long-term orbital evolutions of breakup fragments may conclude that their orbital planes will converge into several corresponding regions in inertial space even if the breakup epoch is not specified. This empirical method combines the aforementioned conclusion with the search strategy developed at Kyushu University, which can identify origins of observed objects as fragments released from a specified spacecraft. This practical method starts with selecting a spacecraft that experienced an orbital anomaly, and formulates a hypothesis to generate fragments from the anomaly. Then, the search strategy is applied to predict the behavior of groups of fragments hypothetically generated. Outcome of this predictive analysis specifies effectively when, where and how we should conduct optical measurements using ground-based telescopes. Objects detected based on the outcome are supposed to be from the anomaly, so that we can confirm the anomaly as a spacecraft breakup to release the detected objects. This paper also demonstrates observation planning for a spacecraft anomaly in the geostationary region.

Orbit Maintenance and Navigation of Human Spacecraft at Cislunar Near Rectilinear Halo Orbits

NASA Technical Reports Server (NTRS)

Davis, Diane; Bhatt, Sagar; Howell, Kathleen; Jang, Jiann-Woei; Whitley, Ryan; Clark, Fred; Guzzetti, Davide; Zimovan, Emily; Barton, Gregg

2017-01-01

Multiple studies have concluded that Earth-Moon libration point orbits are attractive candidates for staging operations. The Near Rectilinear Halo Orbit (NRHO), a member of the Earth-Moon halo orbit family, has been singularly demonstrated to meet multi-mission architectural constraints. In this paper, the challenges associated with operating human spacecraft in the NRHO are evaluated. Navigation accuracies and human vehicle process noise effects are applied to various station keeping strategies in order to obtain a reliable orbit maintenance algorithm. Additionally, the ability to absorb missed burns, construct phasing maneuvers to avoid eclipses and conduct rendezvous and proximity operations are examined.

Coupled orbit-attitude mission design in the circular restricted three-body problem

NASA Astrophysics Data System (ADS)

Guzzetti, Davide

Trajectory design increasingly leverages multi-body dynamical structures that are based on an understanding of various types of orbits in the Circular Restricted Three-Body Problem (CR3BP). Given the more complex dynamical environment, mission applications may also benefit from deeper insight into the attitude motion. In this investigation, the attitude dynamics are coupled with the trajectories in the CR3BP. In a highly sensitive dynamical model, such as the orbit-attitude CR3BP, periodic solutions allow delineation of the fundamental dynamical structures. Periodic solutions are also a subset of motions that are bounded over an infinite time-span (assuming no perturbing factors), without the necessity to integrate over an infinite time interval. Euler equations of motion and quaternion kinematics describe the rotational behavior of the spacecraft, whereas the translation of the center of mass is modeled in the CR3BP equations. A multiple shooting and continuation procedure are employed to target orbit-attitude periodic solutions in this model. Application of Floquet theory, Poincare mapping, and grid search to identify initial guesses for the targeting algorithm is described. In the Earth-Moon system, representative scenarios are explored for axisymmetric vehicles with various inertia characteristics, assuming that the vehicles move along Lyapunov, halo as well as distant retrograde orbits. A rich structure of possible periodic behaviors appears to pervade the solution space in the coupled problem. The stability analysis of the attitude dynamics for the selected families is included. Among the computed solutions, marginally stable and slowly diverging rotational behaviors exist and may offer interesting mission applications. Additionally, the solar radiation pressure is included and a fully coupled orbit-attitude model is developed. With specific application to solar sails, various guidance algorithms are explored to direct the spacecraft along a desired path, when the mutual interaction between orbit and attitude dynamics is considered. Each strategy implements a different form of control input, ranging from instantaneous reorientation of the sail pointing direction to the application of control torques, and it is demonstrated within a simple station keeping scenario.

A demonstration of high precision GPS orbit determination for geodetic applications

NASA Technical Reports Server (NTRS)

Lichten, S. M.; Border, J. S.

1987-01-01

High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.

Designing Delta-DOR acquisition strategies to determine highly elliptical earth orbits

NASA Technical Reports Server (NTRS)

Frauenholz, R. B.

1986-01-01

Delta-DOR acquisition strategies are designed for use in determining highly elliptical earth orbits. The requirements for a possible flight demonstration are evaluated for the Charged Composition Explorer spacecraft of the Active Magnetospheric Particle Tracer Explorers. The best-performing strategy uses data spanning the view periods of two orthogonal baselines near the same orbit periapse. The rapidly changing viewing geometry yields both angular position and velocity information, but each observation may require a different reference quasar. The Delta-DOR data noise is highly dependent on acquisition geometry, varying several orders of magnitude across the baseline view periods. Strategies are selected to minimize the measurement noise predicted by a theoretical model. Although the CCE transponder is limited by S-band and a small bandwidth, the addition of Delta-DOR to coherent Doppler and range improves the one-sigma apogee position accuracy by more than an order of magnitude. Additional Delta-DOR accuracy improvements possible using dual-frequency (S/X) calibration, increased spanned bandwidth, and water-vapor radiometry are presented for comparison. With these benefits, the residual Delta-DOR data noise is primarily due to quasar position uncertainties.

Nuclear cycler: An incremental approach to the deflection of asteroids

NASA Astrophysics Data System (ADS)

Vasile, Massimiliano; Thiry, Nicolas

2016-04-01

This paper introduces a novel deflection approach based on nuclear explosions: the nuclear cycler. The idea is to combine the effectiveness of nuclear explosions with the controllability and redundancy offered by slow push methods within an incremental deflection strategy. The paper will present an extended model for single nuclear stand-off explosions in the proximity of elongated ellipsoidal asteroids, and a family of natural formation orbits that allows the spacecraft to deploy multiple bombs while being shielded by the asteroid during the detonation.

Integrated orbit and attitude hardware-in-the-loop simulations for autonomous satellite formation flying

NASA Astrophysics Data System (ADS)

Park, Han-Earl; Park, Sang-Young; Kim, Sung-Woo; Park, Chandeok

2013-12-01

Development and experiment of an integrated orbit and attitude hardware-in-the-loop (HIL) simulator for autonomous satellite formation flying are presented. The integrated simulator system consists of an orbit HIL simulator for orbit determination and control, and an attitude HIL simulator for attitude determination and control. The integrated simulator involves four processes (orbit determination, orbit control, attitude determination, and attitude control), which interact with each other in the same way as actual flight processes do. Orbit determination is conducted by a relative navigation algorithm using double-difference GPS measurements based on the extended Kalman filter (EKF). Orbit control is performed by a state-dependent Riccati equation (SDRE) technique that is utilized as a nonlinear controller for the formation control problem. Attitude is determined from an attitude heading reference system (AHRS) sensor, and a proportional-derivative (PD) feedback controller is used to control the attitude HIL simulator using three momentum wheel assemblies. Integrated orbit and attitude simulations are performed for a formation reconfiguration scenario. By performing the four processes adequately, the desired formation reconfiguration from a baseline of 500-1000 m was achieved with meter-level position error and millimeter-level relative position navigation. This HIL simulation demonstrates the performance of the integrated HIL simulator and the feasibility of the applied algorithms in a real-time environment. Furthermore, the integrated HIL simulator system developed in the current study can be used as a ground-based testing environment to reproduce possible actual satellite formation operations.

Modal control of an unstable periodic orbit

NASA Astrophysics Data System (ADS)

Wiesel, W.; Shelton, W.

1983-03-01

Floquet theory is applied to the problem of designing a control system for a satellite in an unstable periodic orbit. Expansion about a periodic orbit produces a time-periodic linear system, which is augmented by a time-periodic control term. It is shown that this can be done such that (1) the application of control produces only inertial accelerations, (2) positive real Poincareexponents are shifted into the left half-plane, and (3) the shift of the exponent is linear with control gain. These developments are applied to an unstable orbit near the earth-moon L(3) point pertubed by the sun. Finally, it is shown that the control theory can be extended to include first order perturbations about the periodic orbit without increase in control cost.

Modal control of an unstable periodic orbit

NASA Technical Reports Server (NTRS)

Wiesel, W.; Shelton, W.

1983-01-01

Floquet theory is applied to the problem of designing a control system for a satellite in an unstable periodic orbit. Expansion about a periodic orbit produces a time-periodic linear system, which is augmented by a time-periodic control term. It is shown that this can be done such that (1) the application of control produces only inertial accelerations, (2) positive real Poincareexponents are shifted into the left half-plane, and (3) the shift of the exponent is linear with control gain. These developments are applied to an unstable orbit near the earth-moon L(3) point pertubed by the sun. Finally, it is shown that the control theory can be extended to include first order perturbations about the periodic orbit without increase in control cost.

Real-time maneuver optimization of space-based robots in a dynamic environment: Theory and on-orbit experiments

NASA Astrophysics Data System (ADS)

Chamitoff, Gregory E.; Saenz-Otero, Alvar; Katz, Jacob G.; Ulrich, Steve; Morrell, Benjamin J.; Gibbens, Peter W.

2018-01-01

This paper presents the development of a real-time path-planning optimization approach to controlling the motion of space-based robots. The algorithm is capable of planning three dimensional trajectories for a robot to navigate within complex surroundings that include numerous static and dynamic obstacles, path constraints and performance limitations. The methodology employs a unique transformation that enables rapid generation of feasible solutions for complex geometries, making it suitable for application to real-time operations and dynamic environments. This strategy was implemented on the Synchronized Position Hold Engage Reorient Experimental Satellite (SPHERES) test-bed on the International Space Station (ISS), and experimental testing was conducted onboard the ISS during Expedition 17 by the first author. Lessons learned from the on-orbit tests were used to further refine the algorithm for future implementations.

Inductive crystal field control in layered metal oxides with correlated electrons

DOE PAGES

Balachandran, P. V.; Cammarata, A.; Nelson-Cheeseman, B. B.; ...

2014-07-25

Here, we show that the NiO 6 crystal field energies can be tailored indirectly via heterovalent A cation ordering in layered (La,A) NiO 4 Ruddlesden-Popper (RP) oxides, where A = Sr, Ca, or Ba, using density functional calculations. We leverage as a driving force the electrostatic interactions between charged [LaO] 1+ and neutral [AO] 0 planes to inductively tune the Ni-O bond distortions, without intentional doping or epitaxial strain, altering the correlated d-orbital energies. We use this strategy to design cation ordered LaCaNiO 4 and LaBaNiO 4 with distortions favoring enhanced Ni e g orbital polarization, and find local electronicmore » structure signatures analogous to those in RP La-cuprates, i.e., parent phases of the high-temperature superconducting oxides.« less

Successful Validation of RNA Purification and Quantitative Real-Time PCR Analysis of Gene Expression on the International Space Station

NASA Technical Reports Server (NTRS)

Tran, L.; Parra, Macarena P.; Jung, J.; Boone, T.; Schonfeld, Julie; Almeida, Eduardo

2017-01-01

The NASA Ames WetLab-2 system was developed to offer new on-orbit gene expression analysis capabilities to ISS researchers and can be used to conduct on-orbit RNA isolation and quantitative real time PCR (RT-qPCR) analysis of gene expression from a wide range of biological samples ranging from microbes to mammalian tissues. On orbit validation included three quantitative PCR (qPCR) runs using an E. coli genomic DNA template pre-loaded at three different concentrations. The flight Ct values for the DNA standards showed no statistically significant differences relative to ground controls although there was increased noise in Ct curves, likely due to microgravity-related bubble retention in the optical windows. RNA was successfully purified from both E. coli and mouse liver samples and successfully generated singleplex, duplex and triplex data although with higher standard deviations than ground controls, also likely due to bubbles. Using volunteer science activities, a potential bubble reduction strategy was tested and resulted in smooth amplification curves and tighter Cts between replicates. The WetLab-2 validation experiment demonstrates a novel molecular biology workbench on ISS which allows scientists to purify and stabilize RNA, and to conduct RT-qPCR analyses on-orbit with rapid results. This novel ability is an important step towards utilizing ISS as a National Laboratory facility with the capability to conduct and adjust science experiments in real time without sample return, and opens new possibilities for rapid medical diagnostics and biological environmental monitoring on ISS.

First Results from a Hardware-in-the-Loop Demonstration of Closed-Loop Autonomous Formation Flying

NASA Technical Reports Server (NTRS)

Gill, E.; Naasz, Bo; Ebinuma, T.

2003-01-01

A closed-loop system for the demonstration of autonomous satellite formation flying technologies using hardware-in-the-loop has been developed. Making use of a GPS signal simulator with a dual radio frequency outlet, the system includes two GPS space receivers as well as a powerful onboard navigation processor dedicated to the GPS-based guidance, navigation, and control of a satellite formation in real-time. The closed-loop system allows realistic simulations of autonomous formation flying scenarios, enabling research in the fields of tracking and orbit control strategies for a wide range of applications. The autonomous closed-loop formation acquisition and keeping strategy is based on Lyapunov's direct control method as applied to the standard set of Keplerian elements. This approach not only assures global and asymptotic stability of the control but also maintains valuable physical insight into the applied control vectors. Furthermore, the approach can account for system uncertainties and effectively avoids a computationally expensive solution of the two point boundary problem, which renders the concept particularly attractive for implementation in onboard processors. A guidance law has been developed which strictly separates the relative from the absolute motion, thus avoiding the numerical integration of a target trajectory in the onboard processor. Moreover, upon using precise kinematic relative GPS solutions, a dynamical modeling or filtering is avoided which provides for an efficient implementation of the process on an onboard processor. A sample formation flying scenario has been created aiming at the autonomous transition of a Low Earth Orbit satellite formation from an initial along-track separation of 800 m to a target distance of 100 m. Assuming a low-thrust actuator which may be accommodated on a small satellite, a typical control accuracy of less than 5 m has been achieved which proves the applicability of autonomous formation flying techniques to formations of satellites as close as 50 m.

Near-optimal strategies for sub-decimeter satellite tracking with GPS

NASA Technical Reports Server (NTRS)

Yunck, Thomas P.; Wu, Sien-Chong; Wu, Jiun-Tsong

1986-01-01

Decimeter tracking of low Earth orbiters using differential Global Positioning System (GPS) techniques is discussed. A precisely known global network of GPS ground receivers and a receiver aboard the user satellite are needed, and all techniques simultaneously estimate the user and GPS satellite orbits. Strategies include a purely geometric, a fully dynamic, and a hybrid strategy. The last combines dynamic GPS solutions with a geometric user solution. Two powerful extensions of the hybrid strategy show the most promise. The first uses an optimized synthesis of dynamics and geometry in the user solution, while the second uses a gravity adjustment method to exploit data from repeat ground tracks. These techniques promise to deliver subdecimeter accuracy down to the lowest satellite altitudes.

Ground controlled robotic assembly operations for Space Station Freedom

NASA Technical Reports Server (NTRS)

Parrish, Joseph C.

1991-01-01

A number of dextrous robotic systems and associated positioning and transportation devices are available on Space Station Freedom (SSF) to perform assembly tasks that would otherwise need to be performed by extravehicular activity (EVA) crewmembers. The currently planned operating mode for these robotic systems during the assembly phase is teleoperation by intravehicular activity (IVA) crewmembers. While this operating mode is less hazardous and expensive than manned EVA operations, and has insignificant control loop time delays, the amount of IVA time available to support telerobotic operations is much less than the anticipated requirements. Some alternative is needed to allow the robotic systems to perform useful tasks without exhausting the available IVA resources; ground control is one such alternative. The issues associated with ground control of SSF robotic systems to alleviate onboard crew time availability constraints are investigated. Key technical issues include the effect of communication time delays, the need for safe, reliable execution of remote operations, and required modifications to the SSF ground and flight system architecture. Time delay compensation techniques such as predictive displays and world model-based force reflection are addressed and collision detection and avoidance strategies to ensure the safety of the on-orbit crew, Orbiter, and SSF are described. Although more time consuming and difficult than IVA controlled teleoperations or manned EVA, ground controlled telerobotic operations offer significant benefits during the SSF assembly phase, and should be considered in assembly planning activities.

Enhanced GPS-based GRACE baseline determination by using a new strategy for ambiguity resolution and relative phase center variation corrections

NASA Astrophysics Data System (ADS)

Gu, Defeng; Ju, Bing; Liu, Junhong; Tu, Jia

2017-09-01

Precise relative position determination is a prerequisite for radar interferometry by formation flying satellites. It has been shown that this can be achieved by high-quality, dual-frequency GPS receivers that provide precise carrier-phase observations. The precise baseline determination between satellites flying in formation can significantly improve the accuracy of interferometric products, and has become a research interest. The key technologies of baseline determination using spaceborne dual-frequency GPS for gravity recovery and climate experiment (GRACE) formation are presented, including zero-difference (ZD) reduced dynamic orbit determination, double-difference (DD) reduced dynamic relative orbit determination, integer ambiguity resolution and relative receiver antenna phase center variation (PCV) estimation. We propose an independent baseline determination method based on a new strategy of integer ambiguity resolution and correction of relative receiver antenna PCVs, and implement the method in the NUDTTK software package. The algorithms have been tested using flight data over a period of 120 days from GRACE. With the original strategy of integer ambiguity resolution based on Melbourne-Wübbena (M-W) combinations, the average success rate is 85.6%, and the baseline precision is 1.13 mm. With the new strategy of integer ambiguity resolution based on a priori relative orbit, the average success rate and baseline precision are improved by 5.8% and 0.11 mm respectively. A relative ionosphere-free phase pattern estimation result is given in this study, and with correction of relative receiver antenna PCVs, the baseline precision is further significantly improved by 0.34 mm. For ZD reduced dynamic orbit determination, the orbit precision for each GRACE satellite A or B in three dimensions (3D) is about 2.5 cm compared to Jet Propulsion Laboratory (JPL) post science orbits. For DD reduced dynamic relative orbit determination, the final baseline precision for two GRACE satellites formation is 0.68 mm validated by K-Band Ranging (KBR) observations, and average ambiguity success rate of about 91.4% could be achieved.

Constrained orbital intercept-evasion

NASA Astrophysics Data System (ADS)

Zatezalo, Aleksandar; Stipanovic, Dusan M.; Mehra, Raman K.; Pham, Khanh

2014-06-01

An effective characterization of intercept-evasion confrontations in various space environments and a derivation of corresponding solutions considering a variety of real-world constraints are daunting theoretical and practical challenges. Current and future space-based platforms have to simultaneously operate as components of satellite formations and/or systems and at the same time, have a capability to evade potential collisions with other maneuver constrained space objects. In this article, we formulate and numerically approximate solutions of a Low Earth Orbit (LEO) intercept-maneuver problem in terms of game-theoretic capture-evasion guaranteed strategies. The space intercept-evasion approach is based on Liapunov methodology that has been successfully implemented in a number of air and ground based multi-player multi-goal game/control applications. The corresponding numerical algorithms are derived using computationally efficient and orbital propagator independent methods that are previously developed for Space Situational Awareness (SSA). This game theoretical but at the same time robust and practical approach is demonstrated on a realistic LEO scenario using existing Two Line Element (TLE) sets and Simplified General Perturbation-4 (SGP-4) propagator.

Periodic driving control of Raman-induced spin-orbit coupling in Bose-Einstein condensates: The heating mechanisms

NASA Astrophysics Data System (ADS)

Gomez Llorente, J. M.; Plata, J.

2016-06-01

We focus on a technique recently implemented for controlling the magnitude of synthetic spin-orbit coupling (SOC) in ultracold atoms in the Raman-coupling scenario. This technique uses a periodic modulation of the Raman-coupling amplitude to tune the SOC. Specifically, it has been shown that the effect of a high-frequency sinusoidal modulation of the Raman-laser intensity can be incorporated into the undriven Hamiltonian via effective parameters, whose adiabatic variation can therefore be used to tune the SOC. Here, we characterize the heating mechanisms that can be relevant to this method. We identify the main mechanism responsible for the heating observed in the experiments as basically rooted in driving-induced transfer of population to excited states. Characteristics of that process determined by the harmonic trapping, the decay of the excited states, and the technique used for preparing the system are discussed. Additional heating, rooted in departures from adiabaticity in the variation of the effective parameters, is also described. Our analytical study provides some clues that may be useful in the design of strategies for curbing the effects of heating on the efficiency of the control methods.

Post-aerocapture orbit selection and maintenance for the Aerofast mission to Mars

NASA Astrophysics Data System (ADS)

Pontani, Mauro; Teofilatto, Paolo

2012-10-01

Aerofast is the abbreviation of “aerocapture for future space transportation” and represents a project aimed at developing aerocapture techniques with regard to an interplanetary mission to Mars, in the context of the 7th Framework Program, with the financial support of the European Union. This paper describes the fundamental characteristics of the operational orbit after aerocapture for the mission of interest, as well as the related maintenance strategy. The final orbit selection depends on the desired lighting conditions, maximum revisit time of specific target regions, and feasibility of the orbit maintenance strategy. A sunsynchronous, frozen, repeating-ground-track orbit is chosen. First, the period of repetition is such that adjacent ascending node crossings (over the Mars surface) have a separation compatible with the swath of the optical payload. Secondly, the sunsynchronism condition ensures that a given latitude is periodically visited at the same local time, which condition is essential for comparing images of the same region at different epochs. Lastly, the fulfillment of the frozen condition guarantees improved orbit stability with respect to perturbations due to the zonal harmonics of Mars gravitational field. These three fundamental features of the operational orbit lead to determining its mean orbital elements. The evaluation of short and long period effects (e.g., those due to the sectorial harmonics of the gravitational field or to the aerodynamic drag) requires the determination of the osculating orbital elements at an initial reference time. This research describes a simple and accurate approach that leads to numerically determining these initial values, without employing complicated analytical developments. Numerical simulations demonstrate the long-period stability of the orbit when a significant number of harmonics of the gravitational field are taken into account. However, aerodynamic drag produces a relatively slow orbital decay at the altitudes considered for the mission. This circumstance implies the progressive loss of the sunsynchronism condition, and therefore corrective maneuvers are to be performed. This work proves that actually only in-plane maneuvers are necessary, evaluates the overall delta-v budget needed in the period of repetition (85 Martian nodal days), and proposes a simple maintenance strategy, making reference to the worst-case scenario, which corresponds to the highest seasonal values of the atmospheric density and to the maximum value of the ballistic coefficient of the spacecraft.

An optical survey for space debris on highly eccentric and inclined MEO orbits

NASA Astrophysics Data System (ADS)

Schildknecht, Thomas; Flohrer, Tim; Hinze, Andreas; Vananti, Alessandro; Silha, Jiri

Optical surveys for space debris in high-altitude orbits have been conducted since more than fifteen years. Originally these efforts concentrated mainly on the geostationary ring (GEO) and its close region. Corresponding observation strategies, processing techniques and cataloguing approaches have been developed and successfully applied. The ESA GEO surveys, e.g., resulted in the detection of a significant population of small-size debris and later in the discovery of high area-to-mass ratio objects in GEO-like orbits. The observation scenarios were successively adapted to survey the geostationary transfer orbit (GTO) region; and surveys to search for debris in the medium Earth orbit (MEO) region of the global navigation satellite constellations were successfully conducted. Comparably less experience (both, in terms of practical observation and strategy definition) is available for eccentric orbits that (at least partly) are in the MEO region, in particular for the Molniya-type orbits. Several breakup events and deliberate fragmentations are known to have taken place in such orbits. Survey and follow-up strategies for searching space debris objects in highly-eccentric MEO orbits, and to acquire orbits which are sufficiently accurate to catalogue such objects and to maintain their orbits over longer time spans were developed and, eventually, optical observations were conducted in the framework of an ESA study using ESA' Space Debris Telescope (ESASDT) the 1-m Zeiss telescope located at the Optical Ground Station (OGS) at the Teide Observatory at Tenerife, Spain. Thirteen nights of surveys of Molniya-type orbits was performed between January and August 2013. A basic survey consisted of observing a single geocentric field for 10 minutes. If a faint object was found, follow-up observations were performed during the same night to ensure a save rediscovery of the object during the next nights. Additional follow-up observations to maintain the orbits of these newly discovered faint objects were also acquired with AIUB's 1 m ZIMLAT telescope in Zimmerwald, Switzerland. Eventually, 255 basic surveys were performed during these thirteen nights corresponding to about 47 hours of observations. In total 30 uncorrelated faint objects were discovered. On average one uncorrelated object was found every 100 minutes. Some of these objects show a considerable brightness variation and have a high area-to-mass ratio as determined in the orbit estimation process. We also investigated the detection efficiency of our surveys by comparing the observation results with the TLE population by means of ESA's PROOF tool. Furthermore a comparison of the real detections with the statistical population of the ESA MASTER-2009 model was performed. The result shows that the fragment population of objects in Molniya-type orbits is underestimated in the MASTER model.

The Magnetospheric Multiscale Mission...Resolving Fundamental Processes in Space Plasmas

NASA Technical Reports Server (NTRS)

Curtis, S.

1999-01-01

The Magnetospheric Multiscale (MMS) mission is a multiple-spacecraft Solar-Terrestrial Probe designed to study the microphysics of magnetic reconnection, charged particle acceleration, and turbulence in key boundary regions of Earth's magnetosphere. These three processes, which control the flow of energy, mass, and momentum within and across plasma boundaries, occur throughout the universe and are fundamental to our understanding of astrophysical and solar system plasmas. Only in Earth's magnetosphere, however, are they readily accessible for sustained study through in-situ measurement. MMS will employ five co-orbiting spacecraft identically instrumented to measure electric and magnetic fields, plasmas, and energetic particles. The initial parameters of the individual spacecraft orbits will be designed so that the spacecraft formation will evolve into a three-dimensional configuration near apogee, allowing MMS to differentiate between spatial and temporal effects and to determine the three dimensional geometry of plasma, field, and current structures. In order to sample all of the magnetospheric boundary regions, MMS will employ a unique four-phase orbital strategy involving carefully sequenced changes in the local time and radial distance of apogee and, in the third phase, a change in orbit inclination from 10 degrees to 90 degrees. The nominal mission operational lifetime is two years. Launch is currently scheduled for 2006.

Typical teleoperator time delay profiles, phase 2. [remotely controlled manipulator arms

NASA Technical Reports Server (NTRS)

Wetherington, R. D.; Walsh, J. R.

1974-01-01

The results of the second phase of a study on time delays in communications systems applicable to the teleoperator program are presented. Estimates of the maximum time delays that will be encountered and presents time delay profiles are given for (1) ground control to teleoperator in low earth orbit, (2) ground control to teleoperator in geosynchronous orbit, and (3) low earth orbit control to teleoperator in low earth orbit.

Modeling, simulation and control for a cryogenic fluid management facility, preliminary report

NASA Technical Reports Server (NTRS)

Turner, Max A.; Vanbuskirk, P. D.

1986-01-01

The synthesis of a control system for a cryogenic fluid management facility was studied. The severe demand for reliability as well as instrumentation and control unique to the Space Station environment are prime considerations. Realizing that the effective control system depends heavily on quantitative description of the facility dynamics, a methodology for process identification and parameter estimation is postulated. A block diagram of the associated control system is also produced. Finally, an on-line adaptive control strategy is developed utilizing optimization of the velocity form control parameters (proportional gains, integration and derivative time constants) in appropriate difference equations for direct digital control. Of special concern are the communications, software and hardware supporting interaction between the ground and orbital systems. It is visualized that specialist in the OSI/ISO utilizing the Ada programming language will influence further development, testing and validation of the simplistic models presented here for adaptation to the actual flight environment.

A passive satellite deorbiting strategy for medium earth orbit using solar radiation pressure and the J2 effect

NASA Astrophysics Data System (ADS)

Lücking, Charlotte; Colombo, Camilla; McInnes, Colin R.

2012-08-01

The growing population of space debris poses a serious risk to the future of space flight. To effectively manage the increase of debris in orbit, end-of life disposal has become a key requirement for future missions. This poses a challenge for Medium Earth Orbit (MEO) spacecraft which require a large Δv to re-enter the atmosphere or reach the geostationary graveyard orbit. This paper further explores a passive strategy based on the joint effects of solar radiation pressure and the Earth's oblateness acting on a high area-to-mass-ratio object. The concept was previously presented as an analytical planar model. This paper uses a full 3D model to validate the analytical results numerically for equatorial circular orbits first, then investigating higher inclinations. It is shown that for higher inclinations the initial position of the Sun and right ascension of the ascending node become increasingly important. A region of very low required area-to-mass-ratio is identified in the parameter space of semi-major axis and inclination which occurs for altitudes below 10,000 km.

Orbital Transfer Techniques for Round-Trip Mars Missions

NASA Technical Reports Server (NTRS)

Landau, Damon

2013-01-01

The human exploration of Phobos and Deimos or the retrieval of a surface sample launched to low-Mars orbit presents a highly constrained orbital transfer problem. In general, the plane of the target orbit will not be accessible from the arrival or departure interplanetary trajectories with an (energetically optimal) tangential burn at periapsis. The orbital design is further complicated by the addition of a high-energy parking orbit for the relatively massive Deep Space Vehicle to reduce propellant expenditure, while the crew transfers to and from the target orbit in a smaller Space Exploration Vehicle. The proposed strategy shifts the arrival and departure maneuvers away from periapsis so that the apsidal line of the parking orbit lies in the plane of the target orbit, permitting highly efficient plane change maneuvers at apoapsis of the elliptical parking orbit. An apsidal shift during the arrival or departure maneuver is approximately five times as efficient as maneuvering while in Mars orbit, thus significantly reducing the propellant necessary to transfer between the arrival, target, and departure orbits.

An Orbit And Dispersion Correction Scheme for the PEP II

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

Cai, Y.; Donald, M.; Shoaee, H.

2011-09-01

To achieve optimum luminosity in a storage ring it is vital to control the residual vertical dispersion. In the original PEP storage ring, a scheme to control the residual dispersion function was implemented using the ring orbit as the controlling element. The 'best' orbit not necessarily giving the lowest vertical dispersion. A similar scheme has been implemented in both the on-line control code and in the simulation code LEGO. The method involves finding the response matrices (sensitivity of orbit/dispersion at each Beam-Position-Monitor (BPM) to each orbit corrector) and solving in a least squares sense for minimum orbit, dispersion function ormore » both. The optimum solution is usually a subset of the full least squares solution. A scheme of simultaneously correcting the orbits and dispersion has been implemented in the simulation code and on-line control system for PEP-II. The scheme is based on the eigenvector decomposition method. An important ingredient of the scheme is to choose the optimum eigenvectors that minimize the orbit, dispersion and corrector strength. Simulations indicate this to be a very effective way to control the vertical residual dispersion.« less

Orbital fractures due to domestic violence: an epidemiologic study.

PubMed

Goldberg, Stuart H.; McRill, Connie M.; Bruno, Christopher R.; Ten Have, Tom; Lehman, Erik

2000-09-01

Domestic violence is an important cause of orbital fractures in women. Physicians who treat patients with orbital fractures may not suspect this mechanism of injury. The purpose of this study was to assess the association between domestic violence and orbital fractures. A medical center-based case-control study with matching on age and site of admission was done. Medical center databases were searched using ICD-9 codes to identify all cases of orbital fractures encountered during a three-year period. Medical records of female patients age 13 and older were reviewed along with those of age, gender and site of admission matched controls. A stratified exact test was employed to test the association between domestic violence and orbital fracture. Among 41 adult female cases with orbital fractures treated at our medical center, three (7.3%) reported domestic violence compared to zero among the matched controls (p = 0.037). We believe that domestic violence may be under-reported in both orbital fracture cases and controls. This may result in an underestimate of the orbital fracture versus domestic violence association. Domestic violence is a serious women's health and societal problem. Domestic violence may have a variety of presentations, including illnesses and injuries. Orbital fracture is an identifiable manifestation of domestic violence. Domestic violence is more likely to be detected in adult female hospital patients with orbital fracture than in matched controls with any other diagnosis. Physicians who treat patients with orbital fractures should be familiar with this mechanism of injury.

Preliminary control/structure interaction study of coupled Space Station Freedom/Assembly Work Platform/orbiter

NASA Technical Reports Server (NTRS)

Singh, Sudeep K.; Lindenmoyer, Alan J.

1989-01-01

Results are presented from a preliminary control/structure interaction study of the Space Station, the Assembly Work Platform, and the STS orbiter dynamics coupled with the orbiter and station control systems. The first three Space Station assembly flight configurations and their finite element representations are illustrated. These configurations are compared in terms of control authority in each axis and propellant usage. The control systems design parameters during assembly are computed. Although the rigid body response was acceptable with the orbiter Primary Reaction Control System, the flexible body response showed large structural deflections and loads. It was found that severe control/structure interaction occurred if the stiffness of the Assembly Work Platform was equal to that of the station truss. Also, the response of the orbiter Vernier Reaction Control System to small changes in inertia properties is examined.

Formation Design Strategy for SCOPE High-Elliptic Formation Flying Mission

NASA Technical Reports Server (NTRS)

Tsuda, Yuichi

2007-01-01

The new formation design strategy using simulated annealing (SA) optimization is presented. The SA algorithm is useful to survey a whole solution space of optimum formation, taking into account realistic constraints composed of continuous and discrete functions. It is revealed that this method is not only applicable for circular orbit, but also for high-elliptic orbit formation flying. The developed algorithm is first tested with a simple cart-wheel motion example, and then applied to the formation design for SCOPE. SCOPE is the next generation geomagnetotail observation mission planned in JAXA, utilizing a formation flying techonology in a high elliptic orbit. A distinctive and useful heuristics is found by investigating SA results, showing the effectiveness of the proposed design process.

Applicability of the control configured design approach to advanced earth orbital transportation systems

NASA Technical Reports Server (NTRS)

Hepler, A. K.; Zeck, H.; Walker, W. H.; Shafer, D. E.

1978-01-01

The applicability of the control configured design approach (CCV) to advanced earth orbital transportation systems was studied. The baseline system investigated was fully reusable vertical take-off/horizontal landing single-stage-to-orbit vehicle and had mission requirements similar to the space shuttle orbiter. Technical analyses were made to determine aerodynamic, flight control and subsystem design characteristics. Figures of merit were assessed on vehicle dry weight and orbital payload. The results indicated that the major parameters for CCV designs are hypersonic trim, aft center of gravity, and control surface heating. Optimized CCV designs can be controllable and provide substantial payload gains over conventional non-CCV design vertical take-off vehicles.

Development of a Two-Wheel Contingency Mode for the MAP Spacecraft

NASA Technical Reports Server (NTRS)

Starin, Scott R.; ODonnell, James R., Jr.; Bauer, Frank H. (Technical Monitor)

2002-01-01

In the event of a failure of one of MAP's three reaction wheel assemblies (RWAs), it is not possible to achieve three-axis, full-state attitude control using the remaining two wheels. Hence, two of the attitude control algorithms implemented on the MAP spacecraft will no longer be usable in their current forms: Inertial Mode, used for slewing to and holding inertial attitudes, and Observing Mode, which implements the nominal dual-spin science mode. This paper describes the effort to create a complete strategy for using software algorithms to cope with a RWA failure. The discussion of the design process will be divided into three main subtopics: performing orbit maneuvers to reach and maintain an orbit about the second Earth-Sun libration point in the event of a RWA failure, completing the mission using a momentum-bias two-wheel science mode, and developing a new thruster-based mode for adjusting the inertially fixed momentum bias. In this summary, the philosophies used in designing these changes is shown; the full paper will supplement these with algorithm descriptions and testing results.

Real-time control of optimal low-thrust transfer to the Sun-Earth L 1 halo orbit in the bicircular four-body problem

NASA Astrophysics Data System (ADS)

Salmani, Majid; Büskens, Christof

2011-11-01

In this article, after describing a procedure to construct trajectories for a spacecraft in the four-body model, a method to correct the trajectory violations is presented. To construct the trajectories, periodic orbits as the solutions of the three-body problem are used. On the other hand, the bicircular model based on the Sun-Earth rotating frame governs the dynamics of the spacecraft and other bodies. A periodic orbit around the first libration-point L1 is the destination of the mission which is one of the equilibrium points in the Sun-Earth/Moon three-body problem. In the way to reach such a far destination, there are a lot of disturbances such as solar radiation and winds that make the plans untrustworthy. However, the solar radiation pressure is considered in the system dynamics. To prevail over these difficulties, considering the whole transfer problem as an optimal control problem makes the designer to be able to correct the unavoidable violations from the pre-designed trajectory and strategies. The optimal control problem is solved by a direct method, transcribing it into a nonlinear programming problem. This transcription gives an unperturbed optimal trajectory and its sensitivities with respect perturbations. Modeling these perturbations as parameters embedded in a parametric optimal control problem, one can take advantage of the parametric sensitivity analysis of nonlinear programming problem to recalculate the optimal trajectory with a very smaller amount of computation costs. This is obtained by evaluating a first-order Taylor expansion of the perturbed solution in an iterative process which is aimed to achieve an admissible solution. At the end, the numerical results show the applicability of the presented method.

Minimum dV for Targeted Spacecraft Disposal

NASA Technical Reports Server (NTRS)

Bacon, John

2017-01-01

The density scale height of the Earth's atmosphere undergoes significant reduction in the final phases of a natural decay. It can be shown that for most realistic ballistic numbers, it is possible to exploit this effect to amplify available spacecraft dV by using it at the penultimate perigee to penetrate into higher drag regions at final apogee. The drag at this lower pass can more effectively propel a spacecraft towards the final target region than applying the same dV direct Hohmann transfer at that final apogee. This study analyzes the potential use of this effect-- in combination with small phasing burns--to calculate the absolute minimum delta-V that would be required to reliably guide a spacecraft to any specified safe unoccupied ocean region as a function of ballistic number, orbit inclination, and initial eccentricity. This calculation is made for controllable spacecraft in several orbit inclinations and eccentricities with arbitrary initial LAN and ArgP one week before final entry, under three-sigma atmospheric perturbations. The study analyzes the dV required under varying levels of final controllable altitude at which dV may be imparted, and various definitions of the length and location of a "safe" disposal area. The goal of such research is to improve public safety by creating assured safe disposal strategies for low-dV and/or low-thrust spacecraft that under more traditional strategies would need to be abandoned to a fully random decay.

Detail view of the "underside" of the Orbiter Maneuvering/Reaction Control ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of the "underside" of the Orbiter Maneuvering/Reaction Control Systems pod looking at the two spherical propellant tanks for the Reaction Control System, and the elongated propellant tanks for the Orbiter Maneuvering System. This view was taken at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Periodic orbits of solar sail equipped with reflectance control device in Earth-Moon system

NASA Astrophysics Data System (ADS)

Yuan, Jianping; Gao, Chen; Zhang, Junhua

2018-02-01

In this paper, families of Lyapunov and halo orbits are presented with a solar sail equipped with a reflectance control device in the Earth-Moon system. System dynamical model is established considering solar sail acceleration, and four solar sail steering laws and two initial Sun-sail configurations are introduced. The initial natural periodic orbits with suitable periods are firstly identified. Subsequently, families of solar sail Lyapunov and halo orbits around the L1 and L2 points are designed with fixed solar sail characteristic acceleration and varying reflectivity rate and pitching angle by the combination of the modified differential correction method and continuation approach. The linear stabilities of solar sail periodic orbits are investigated, and a nonlinear sliding model controller is designed for station keeping. In addition, orbit transfer between the same family of solar sail orbits is investigated preliminarily to showcase reflectance control device solar sail maneuver capability.

Detail view of a starboard Orbiter Maneuvering and Reaction Control ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of a starboard Orbiter Maneuvering and Reaction Control Systems pod, removed from the orbiter and in it's carrier/transport vehicle at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Low-energy near Earth asteroid capture using Earth flybys and aerobraking

NASA Astrophysics Data System (ADS)

Tan, Minghu; McInnes, Colin; Ceriotti, Matteo

2018-04-01

Since the Sun-Earth libration points L1 and L2 are regarded as ideal locations for space science missions and candidate gateways for future crewed interplanetary missions, capturing near-Earth asteroids (NEAs) around the Sun-Earth L1/L2 points has generated significant interest. Therefore, this paper proposes the concept of coupling together a flyby of the Earth and then capturing small NEAs onto Sun-Earth L1/L2 periodic orbits. In this capture strategy, the Sun-Earth circular restricted three-body problem (CRTBP) is used to calculate target Lypaunov orbits and their invariant manifolds. A periapsis map is then employed to determine the required perigee of the Earth flyby. Moreover, depending on the perigee distance of the flyby, Earth flybys with and without aerobraking are investigated to design a transfer trajectory capturing a small NEA from its initial orbit to the stable manifolds associated with Sun-Earth L1/L2 periodic orbits. Finally, a global optimization is carried out, based on a detailed design procedure for NEA capture using an Earth flyby. Results show that the NEA capture strategies using an Earth flyby with and without aerobraking both have the potential to be of lower cost in terms of energy requirements than a direct NEA capture strategy without the Earth flyby. Moreover, NEA capture with an Earth flyby also has the potential for a shorter flight time compared to the NEA capture strategy without the Earth flyby.

Effectiveness of GNSS disposal strategies

NASA Astrophysics Data System (ADS)

Alessi, E. M.; Rossi, A.; Valsecchi, G. B.; Anselmo, L.; Pardini, C.; Colombo, C.; Lewis, H. G.; Daquin, J.; Deleflie, F.; Vasile, M.; Zuiani, F.; Merz, K.

2014-06-01

The management of the Global Navigation Satellite Systems (GNSS) and of the Medium Earth Orbit (MEO) region as a whole is a subject that cannot be deferred, due to the growing exploitation and launch rate in that orbital regime. The advent of the European Galileo and the Chinese Beidou constellations significantly added complexity to the system and calls for an adequate global view on the four constellations present in operation. The operation procedures, including maintenance and disposal practices, of the constellations currently deployed were analyzed in order to asses a proper reference simulation scenario. The complex dynamics of the MEO region with all the geopotential and lunisolar resonances was studied to better identify the proper end-of-life orbit for every proposed strategy, taking into account and, whenever possible, exploiting the orbital dynamics in this peculiar region of space. The possibility to exploit low thrust propulsion or non gravitational perturbations with passive de-orbiting devices (and a combination of the two) was analyzed, in view of possible applications in the design of the future generations of the constellations satellites. Several upgrades in the long-term evolution software SDM and DAMAGE were undertaken to properly handle the constellation simulations in every aspect from constellation maintenance to orbital dynamics. A thorough approach considering the full time evolving covariance matrix associated with every object was implemented in SDM to compute the collision risk and associated maneuver rate for the constellation satellites. Once the software upgrades will be completed, the effectiveness of the different disposal strategies will be analyzed in terms of residual collision risk and avoidance maneuvers rate. This work was performed under the ESA/GSP Contract no. 4000107201/12/F/MOS.

Flight controller alertness and performance during spaceflight shiftwork operations.

PubMed

Kelly, S M; Rosekind, M R; Dinges, D F; Miller, D L; Gillen, K A; Gregory, K B; Aguilar, R D; Smith, R M

1998-09-01

Decreased alertness and performance associated with fatigue, sleep loss, and circadian disruption are issues faced by a diverse range of shiftwork operations personnel. During Space Transportation System (STS) operations, Mission Operations Directorate (MOD) personnel provide 24-hr. coverage of critical tasks. A joint NASA Johnson Space Center and NASA Ames Research Center project was undertaken to examine these issues in flight controllers during MOD shiftwork operations. An initial operational test of procedures and measures was conducted during the STS-53 mission in December 1992. The study measures included a Background Questionnaire, a subjective daily logbook completed on a 24-hour basis (to report sleep patterns, work periods, etc.), and an 8 minute performance and mood test battery administered at the beginning, middle, and end of each shift period. Seventeen flight controllers representing the 3 Orbit shifts participated. The initial results clearly support the need for further data collection during other STS missions to document baseline levels of alertness and performance during MOD shiftwork operations. Countermeasure strategies specific to the MOD environment are being developed to minimize the adverse effects of fatigue, sleep loss, and circadian disruption engendered by shiftwork operations. These issues are especially pertinent for the night shift operations and the acute phase advance required for the transition of day shift personnel into the night for shuttle launch. Implementation and evaluation of the countermeasure strategies to maximize alertness and performance is planned. As STS missions extend to further EDO (extended duration orbiters), and timelines and planning for 24-hour Space Station operations continue, alertness and performance issues related to sleep and circadian disruption will remain highly relevant in the MOD environment.

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

NASA Technical Reports Server (NTRS)

Wong, Hong; Kapila, Vikram

2004-01-01

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

CERES: An ab initio code dedicated to the calculation of the electronic structure and magnetic properties of lanthanide complexes.

PubMed

Calvello, Simone; Piccardo, Matteo; Rao, Shashank Vittal; Soncini, Alessandro

2018-03-05

We have developed and implemented a new ab initio code, Ceres (Computational Emulator of Rare Earth Systems), completely written in C++11, which is dedicated to the efficient calculation of the electronic structure and magnetic properties of the crystal field states arising from the splitting of the ground state spin-orbit multiplet in lanthanide complexes. The new code gains efficiency via an optimized implementation of a direct configurational averaged Hartree-Fock (CAHF) algorithm for the determination of 4f quasi-atomic active orbitals common to all multi-electron spin manifolds contributing to the ground spin-orbit multiplet of the lanthanide ion. The new CAHF implementation is based on quasi-Newton convergence acceleration techniques coupled to an efficient library for the direct evaluation of molecular integrals, and problem-specific density matrix guess strategies. After describing the main features of the new code, we compare its efficiency with the current state-of-the-art ab initio strategy to determine crystal field levels and properties, and show that our methodology, as implemented in Ceres, represents a more time-efficient computational strategy for the evaluation of the magnetic properties of lanthanide complexes, also allowing a full representation of non-perturbative spin-orbit coupling effects. © 2017 Wiley Periodicals, Inc. © 2017 Wiley Periodicals, Inc.

Lessons Learned from the Hubble Space Telescope (HST) Contamination Control Program

NASA Technical Reports Server (NTRS)

Hansen, Patricia A.; Townsend, Jacqueline A.; Hedgeland, Randy J.

2004-01-01

Over the past two decades, the Hubble Space Telescope (HST) Contamination Control Program has evolved from a ground-based integration program to a space-based science-sustaining program. The contamination controls from the new-generation Scientific Instruments and Orbital Replacement Units were incorporated into the HST Contamination Control Program to maintain scientific capability over the life of the telescope. Long-term on-orbit scientific data has shown that these contamination controls implemented for the instruments, Servicing Mission activities (Orbiter, Astronauts, and mission), and on-orbit operations successfully protected the HST &om contamination and the instruments from self-contamination.

Lessons Learned from the Hubble Space Telescope (HST) Contamination Control Program

NASA Technical Reports Server (NTRS)

Hansen, Patricia A.; Townsend, Jacqueline A.; Hedgeland, Randy J.

2004-01-01

Over the past two decades, the Hubble Space Telescope (HST) Contamination Control Program has evolved from a ground-based integration program to a space-based science-sustaining program. The contamination controls from the new-generation Scientific Instruments and Orbital Replacement Units were incorporated into the HST Contamination Control Program to maintain scientific capability over the life of the telescope. Long-term on-orbit scientific data has shown that these contamination controls implemented for the instruments, Servicing Mission activities (Orbiter, Astronauts, and mission), and on-orbit operations successfully protected the HST from contamination and the instruments from self-contamination.

Spaceflight mechanics 1992; Proceedings of the 2nd AAS/AIAA Meeting, Colorado Springs, CO, Feb. 24-26, 1992. Pts. 1 & 2

NASA Astrophysics Data System (ADS)

Diehl, Roger E.; Schinnerer, Ralph G.; Williamson, Walton E.; Boden, Daryl G.

The present conference discusses topics in orbit determination, tethered satellite systems, celestial mechanics, guidance optimization, flexible body dynamics and control, attitude dynamics and control, Mars mission analyses, earth-orbiting mission analysis/debris, space probe mission analyses, and orbital computation numerical analyses. Attention is given to electrodynamic forces for control of tethered satellite systems, orbiting debris threats to asteroid flyby missions, launch velocity requirements for interceptors of short range ballistic missiles, transfers between libration-point orbits in the elliptic restricted problem, minimum fuel spacecraft reorientation, orbital guidance for hitting a fixed point at maximum speed, efficient computation of satellite visibility periods, orbit decay and reentry prediction for space debris, and the determination of satellite close approaches.

Spaceflight mechanics 1992; Proceedings of the 2nd AAS/AIAA Meeting, Colorado Springs, CO, Feb. 24-26, 1992. Pts. 1 & 2

NASA Technical Reports Server (NTRS)

Diehl, Roger E. (Editor); Schinnerer, Ralph G. (Editor); Williamson, Walton E. (Editor); Boden, Daryl G. (Editor)

1992-01-01

The present conference discusses topics in orbit determination, tethered satellite systems, celestial mechanics, guidance optimization, flexible body dynamics and control, attitude dynamics and control, Mars mission analyses, earth-orbiting mission analysis/debris, space probe mission analyses, and orbital computation numerical analyses. Attention is given to electrodynamic forces for control of tethered satellite systems, orbiting debris threats to asteroid flyby missions, launch velocity requirements for interceptors of short range ballistic missiles, transfers between libration-point orbits in the elliptic restricted problem, minimum fuel spacecraft reorientation, orbital guidance for hitting a fixed point at maximum speed, efficient computation of satellite visibility periods, orbit decay and reentry prediction for space debris, and the determination of satellite close approaches.

Correlation of the Hubble Space Telescope (HST) Space Telescope Imaging Spectrometer (STIS) On-Orbit Data with Pre-launch Predictions and Ground Contamination Controls

NASA Technical Reports Server (NTRS)

Hansen, Patricia A.

2003-01-01

The Hubble Space Telescope (HST) Space Telescope Imaging Spectrograph (STIS) was deployed on-orbit in February 1997. The contamination program for STIS was stringently controlled as the five-year end-of-life deposition was set at 158, per optical element. Contamination was controlled through materials selection, extensive vacuum outgassing certifications, cleaning techniques, and environmental controls. In addition to ground contamination controls, on-orbit contamination controls were implemented for both the HST servicing mission activities and early post-servicing mission checkout. The extensive contamination control program will be discussed and the STIS on-orbit data will be correlated with the prelaunch analytical predictions.

Magnetic attitude control torque generation of a gravity gradient stabilized satellite

NASA Astrophysics Data System (ADS)

Suhadis, N. M.; Salleh, M. B.; Rajendran, P.

2018-05-01

Magnetic torquer is used to generate a magnetic dipole moment onboard satellites whereby a control torque for attitude control purposes is generated when it couples with the geomagnetic field. This technique has been considered very attractive for satellites operated in Low Earth Orbit (LEO) as the strength of the geomagnetic field is relatively high below the altitude of 1000 km. This paper presents the algorithm used to generate required magnetic dipole moment by 3 magnetic torquers mounted onboard a gravity gradient stabilized satellite operated at an altitude of 540 km with nadir pointing mission. As the geomagnetic field cannot be altered and its magnitude and direction vary with respect to the orbit altitude and inclination, a comparison study of attitude control torque generation performance with various orbit inclination is performed where the structured control algorithm is simulated for 13°, 33° and 53° orbit inclinations to see how the variation of the satellite orbit affects the satellite's attitude control torque generation. Results from simulation show that the higher orbit inclination generates optimum magnetic attitude control torque for accurate nadir pointing mission.

Smart-1 Moon Impact Operations

NASA Technical Reports Server (NTRS)

Ayala, Andres; Rigger, Ralf

2007-01-01

This paper describes the operations to control the Moon impact of the 3-axis stabilized spacecraft SMART-1 in September 2006. SMART-1 was launched on 27/09/2003. It was the first ESA mission to use an Electric Propulsion (EP) engine as the main motor to spiral out of the Earth gravity field and reach a scientific moon orbit [1]. During September 2005 the last EP maneuvers were performed using the remaining Xenon, in order to compensate for the 3rd body perturbations of the Sun and Earth. These operations extended the mission for an additional year. Afterwards the EP performance became unpredictable and low, so that no meaningful operation for the moon impact could be done. To move the predicted impact point on the 16/8/2006 into visibility from Earth an alternative Delta-V strategy was designed. Due to their alignment, the attitude thrusters could not be used directly to generate the Delta-V, so this strategy was based on controlled angular momentum biasing. Firing along the velocity vector around apolune, the remaining Hydrazine left from the attitude control budget was used, to shift the impact to the required coordinates.

The globe and orbit in Laron syndrome.

PubMed

Kornreich, L; Konen, O; Lilos, P; Laron, Z

2011-09-01

Patients with LS have an inborn growth hormone resistance, resulting in failure to generate IGF-1. The purpose of this study was to evaluate the size of the eye and orbit in LS. We retrospectively reviewed the MR imaging of the brain in 9 patients with LS for the following parameters: axial diameter of the globe, interzygomatic distance, perpendicular distance from the interzygomatic line to margins of the globe, medial-to-lateral diameter of the orbit at the anterior orbital rim, distance from the anterior orbital rim to the anterior globe, maximal distance between the medial walls of the orbits, lateral orbital wall angle, lateral orbital wall length, and mediolateral thickness of the intraorbital fat in the most cranial image of the orbit. All measurements were made bilaterally. Twenty patients referred for MR imaging for unrelated reasons served as control subjects. Compared with the control group, the patients with LS had a significantly smaller maximal globe diameter and shallower but wider orbits due to a shorter lateral wall, a smaller medial distance between the orbits, and a larger angle of the orbit. The ratio between the most anterior orbital diameter and the globe was greater than that in controls. The position of the globe was more anterior in relation to the interzygomatic line. Shallow and wide orbits and small globes relative to orbital size are seen in LS and may be secondary to IGF-1 deficiency.

View of Mission Control during Apollo 9 earth orbital mission

NASA Image and Video Library

1969-03-03

S69-26301 (March 1969) --- Overall view of the Mission Operations Control Room in the Mission Control Center, Building 30, during the Apollo 9 Earth-orbital mission. When this photograph was taken a live television transmission was being received from Apollo 9 as it orbited Earth.

Outcomes in Diabetic Patients Undergoing Orbital Atherectomy System.

PubMed

Lee, Michael S; Shlofmitz, Evan; Nguyen, Heajung; Shlofmitz, Richard A

2016-10-01

We evaluated the angiographic and clinical outcomes of orbital atherectomy to treat severely calcified coronary lesions in diabetic and non-diabetic patients. Diabetics have increased risk for death, myocardial infarction, and target vessel revascularization after percutaneous coronary intervention. Severely calcified coronary lesions are associated with increased cardiac events. Orbital atherectomy facilitates stent delivery and optimizes stent expansion by modifying severely calcified plaque. Outcomes in diabetic patients who undergo orbital atherectomy have not been reported. Our retrospective multicenter registry included 458 consecutive real-world patients with severely calcified coronary arteries who underwent orbital atherectomy. The primary safety endpoint was the rate of major adverse cardiac and cerebrovascular events at 30 days. Diabetics represented 42.1% (193/458) of the entire cohort. The primary endpoint was similar in diabetics and non-diabetics (1.0% vs. 3.0%%, P = 0.20), as were 30-day rates of death (0.5% vs. 1.9%, P = 0.41), myocardial infarction (0.5% vs. 1.5%, P = 0.40), target vessel revascularization (0% vs. 0%, P = 1), and stroke (0% vs. 0.4%, P > 0.9). Angiographic complications and stent thrombosis rate were low and did not differ between the 2 groups. Diabetics represented a sizeable portion of patients who underwent orbital atherectomy. Diabetics who had severely calcified coronary arteries and underwent orbital atherectomy had low event rates that were similar to non-diabetics. Orbital atherectomy appears to be a viable treatment strategy for diabetic patients. Randomized trials with longer-term follow-up are needed to determine the ideal treatment strategy for diabetics. © 2016, Wiley Periodicals, Inc.

Time-Varying Expression of the Formation Flying along Circular Trajectories

NASA Technical Reports Server (NTRS)

Kawaguchi, Jun'ichiro

2007-01-01

Usually, the formation flying associated with circular orbits is discussed through the well-known Hill s or C-W equations of motion. This paper dares to present and discuss the coordinates that may contain time-varying coefficients. The discussion presents how the controller s performance is affected by the selection of coordinates, and also looks at the special coordinate suitable for designating a target bin to which each spacecraft in the formation has only to be guided. It is revealed that the latter strategy may incorporate the J2 disturbance automatically.

Computing Satellite Maneuvers For A Repeating Ground Track

NASA Technical Reports Server (NTRS)

Shapiro, Bruce

1994-01-01

TOPEX/POSEIDON Ground Track Maintenance Maneuver Targeting Program (GTARG) assists in designing maneuvers to maintain orbit of TOPEX/POSEIDON satellite. Targeting strategies used either maximize time between maneuvers or force control band exit to occur at specified intervals. Runout mode allows for ground-track propagation without targeting. GTARG incorporates analytic mean-element propagation algorithm accounting for all perturbations known to cause significant variations in ground track. Perturbations include oblateness of Earth, luni-solar gravitation, drag, thrusts associated with impulsive maneuvers, and unspecified fixed forces acting on satellite in direction along trajectory. Written in VAX-FORTRAN.

Mariner 9 navigation

NASA Technical Reports Server (NTRS)

Neil, W. J.; Jordan, J. F.; Zielenbach, J. W.; Wong, S. K.; Mitchell, R. T.; Webb, W. A.; Koskela, P. E.

1973-01-01

A final, comprehensive description of the navigation of Mariner 9-the first U.S. spacecraft to orbit another planet is provided. The Mariner 9 navigation function included not only precision flight path control but also pointing of the spacecraft's scientific instruments mounted on a two degree of freedom scan platform. To the extent appropriate, each section describes the perflight analyses on which the operational strategies and performance predictions were based. Inflight results are then discussed and compared with the preflight predictions. Postflight analyses, which were primarily concerned with developing a thorough understanding of unexpected in-flight results, are also presented.

Orbit determination error analysis and comparison of station-keeping costs for Lissajous and halo-type libration point orbits and sensitivity analysis using experimental design techniques

NASA Technical Reports Server (NTRS)

Gordon, Steven C.

1993-01-01

Spacecraft in orbit near libration point L1 in the Sun-Earth system are excellent platforms for research concerning solar effects on the terrestrial environment. One spacecraft mission launched in 1978 used an L1 orbit for nearly 4 years, and future L1 orbital missions are also being planned. Orbit determination and station-keeping are, however, required for these orbits. In particular, orbit determination error analysis may be used to compute the state uncertainty after a predetermined tracking period; the predicted state uncertainty levels then will impact the control costs computed in station-keeping simulations. Error sources, such as solar radiation pressure and planetary mass uncertainties, are also incorporated. For future missions, there may be some flexibility in the type and size of the spacecraft's nominal trajectory, but different orbits may produce varying error analysis and station-keeping results. The nominal path, for instance, can be (nearly) periodic or distinctly quasi-periodic. A periodic 'halo' orbit may be constructed to be significantly larger than a quasi-periodic 'Lissajous' path; both may meet mission requirements, but perhaps the required control costs for these orbits are probably different. Also for this spacecraft tracking and control simulation problem, experimental design methods can be used to determine the most significant uncertainties. That is, these methods can determine the error sources in the tracking and control problem that most impact the control cost (output); it also produces an equation that gives the approximate functional relationship between the error inputs and the output.

Optimal spacecraft formation establishment and reconfiguration propelled by the geomagnetic Lorentz force

NASA Astrophysics Data System (ADS)

Huang, Xu; Yan, Ye; Zhou, Yang

2014-12-01

The Lorentz force acting on an electrostatically charged spacecraft as it moves through the planetary magnetic field could be utilized as propellantless electromagnetic propulsion for orbital maneuvering, such as spacecraft formation establishment and formation reconfiguration. By assuming that the Earth's magnetic field could be modeled as a tilted dipole located at the center of Earth that corotates with Earth, a dynamical model that describes the relative orbital motion of Lorentz spacecraft is developed. Based on the proposed dynamical model, the energy-optimal open-loop trajectories of control inputs, namely, the required specific charges of Lorentz spacecraft, for Lorentz-propelled spacecraft formation establishment or reconfiguration problems with both fixed and free final conditions constraints are derived via Gauss pseudospectral method. The effect of the magnetic dipole tilt angle on the optimal control inputs and the relative transfer trajectories for formation establishment or reconfiguration is also investigated by comparisons with the results derived from a nontilted dipole model. Furthermore, a closed-loop integral sliding mode controller is designed to guarantee the trajectory tracking in the presence of external disturbances and modeling errors. The stability of the closed-loop system is proved by a Lyapunov-based approach. Numerical simulations are presented to verify the validity of the proposed open-loop control methods and demonstrate the performance of the closed-loop controller. Also, the results indicate the dipole tilt angle should be considered when designing control strategies for Lorentz-propelled spacecraft formation establishment or reconfiguration.

On-orbit checkout of satellites, volume 2. Part 3 of on-orbit checkout study. [space maintenance

NASA Technical Reports Server (NTRS)

Pritchard, E. I.

1978-01-01

Early satellite failures significantly degrading satellite operations are reviewed with emphasis on LANDSAT D, the Technology Demonstration Satellite, the ATREX/AEM spacecraft, STORMSAT 2, and the synchronous meteorological satellite. Candidates for correction with on-orbit checkout and appropriate actions are analyzed. On-orbit checkout subsystem level studies are summarized for electrical power, attitude control, thermal control, reaction control and propulsion, instruments, and angular rate matching for alignment of satellite IRU.

Optimal mission planning of GEO on-orbit refueling in mixed strategy

NASA Astrophysics Data System (ADS)

Chen, Xiao-qian; Yu, Jing

2017-04-01

The mission planning of GEO on-orbit refueling (OOR) in Mixed strategy is studied in this paper. Specifically, one SSc will be launched to an orbital slot near the depot when multiple GEO satellites are reaching their end of lives. The SSc replenishes fuel from the depot and then extends the lifespan of the target satellites via refueling. In the mixed scenario, only some of the target satellites could be served by the SSc, and the remaining ones will be fueled by Pseudo SScs (the target satellite which has already been refueled by the SSc and now has sufficient fuel for its operation as well as the fuel to refuel other target satellites is called Pseudo SSc here). The mission sequences and fuel mass of the SSc and Pseudo SScs, the dry mass of the SSc are used as design variables, whereas the economic benefit of the whole mission is used as design objective. The economic cost and benefit models are stated first, and then a mathematical optimization model is proposed. A comprehensive solution method involving enumeration, particle swarm optimization and modification is developed. Numerical examples are carried out to demonstrate the effectiveness of the model and solution method. Economic efficiencies of different OOR strategies are compared and discussed. The mixed strategy would perform better than the other strategies only when the target satellites satisfy some conditions. This paper presents an available mixed strategy scheme for users and analyzes its advantages and disadvantages by comparing with some other OOR strategies, providing helpful references to decision makers. The best strategy in practical applications depends on the specific demands and user preference.

Scheduler for monitoring objects orbiting earth using satellite-based telescopes

DOEpatents

Olivier, Scot S; Pertica, Alexander J; Riot, Vincent J; De Vries, Willem H; Bauman, Brian J; Nikolaev, Sergei; Henderson, John R; Phillion, Donald W

2015-04-28

An ephemeris refinement system includes satellites with imaging devices in earth orbit to make observations of space-based objects ("target objects") and a ground-based controller that controls the scheduling of the satellites to make the observations of the target objects and refines orbital models of the target objects. The ground-based controller determines when the target objects of interest will be near enough to a satellite for that satellite to collect an image of the target object based on an initial orbital model for the target objects. The ground-based controller directs the schedules to be uploaded to the satellites, and the satellites make observations as scheduled and download the observations to the ground-based controller. The ground-based controller then refines the initial orbital models of the target objects based on the locations of the target objects that are derived from the observations.

Monitoring objects orbiting earth using satellite-based telescopes

DOEpatents

Olivier, Scot S.; Pertica, Alexander J.; Riot, Vincent J.; De Vries, Willem H.; Bauman, Brian J.; Nikolaev, Sergei; Henderson, John R.; Phillion, Donald W.

2015-06-30

An ephemeris refinement system includes satellites with imaging devices in earth orbit to make observations of space-based objects ("target objects") and a ground-based controller that controls the scheduling of the satellites to make the observations of the target objects and refines orbital models of the target objects. The ground-based controller determines when the target objects of interest will be near enough to a satellite for that satellite to collect an image of the target object based on an initial orbital model for the target objects. The ground-based controller directs the schedules to be uploaded to the satellites, and the satellites make observations as scheduled and download the observations to the ground-based controller. The ground-based controller then refines the initial orbital models of the target objects based on the locations of the target objects that are derived from the observations.

Orbiter ECLSS support of Shuttle payloads

NASA Technical Reports Server (NTRS)

Jaax, J. R.; Morris, D. W.; Prince, R. N.

1974-01-01

The orbiter ECLSS (Environmental Control and Life Support System) provides the functions of atmosphere revitalization, crew life support, and active thermal control. This paper describes these functions as they relate to the support of Shuttle payloads, including automated spacecraft, Spacelab and Department of Defense missions. Functional and performance requirements for the orbiter ECLSS which affect payload support are presented for the atmosphere revitalization subsystem, the food, water and waste subsystem, and the active thermal control subsystem. Schematics for these subsystems are also described. Finally, based on the selected orbiter configuration, preliminary design and off-design thermodynamic data are presented to quantify the baseline orbiter capability; to quantify the payload chargeable penalties for increasing this support; and to identify the significant limits of orbiter ECLSS support available to Shuttle payloads.

Full Stark control of polariton states on a spin-orbit hypersphere

NASA Astrophysics Data System (ADS)

Li, Feng; Cancellieri, E.; Buonaiuto, G.; Skolnick, M. S.; Krizhanovskii, D. N.; Whittaker, D. M.

2016-11-01

The orbital angular momentum and the polarization of light are physical quantities widely investigated for classical and quantum information processing. In this work we propose to take advantage of strong light-matter coupling, circular-symmetric confinement, and transverse-electric transverse-magnetic splitting to exploit states where these two degrees of freedom are combined. To this end we develop a model based on a spin-orbit Poincaré hypersphere. Then we consider the example of semiconductor polariton systems and demonstrate full ultrafast Stark control of spin-orbit states. Moreover, by controlling states on three different spin-orbit spheres and switching from one sphere to another we demonstrate the control of different logic bits within one single physical system.

Two designs for an orbital transfer vehicle

NASA Technical Reports Server (NTRS)

Davis, Richard; Duquette, Miles; Fredrick, Rebecca; Schumacher, Daniel; Somers, Schaeffer; Stafira, Stanley; Williams, James; Zelinka, Mark

1988-01-01

The Orbital Transfer Vehicle (OTV) and systems were researched in the following areas: avionics, crew systems, electrical power systems, environmental control/life support systems, navigation and orbital maneuvers, propulsion systems, reaction control systems (RCS), servicing systems, and structures.

Flight controller alertness and performance during MOD shiftwork operations

NASA Technical Reports Server (NTRS)

Kelly, Sean M.; Rosekind, Mark R.; Dinges, David F.; Miller, Donna L.; Gillen, Kelly A.; Gregory, Kevin B.; Aguilar, Ronald D.; Smith, Roy M.

1994-01-01

Decreased alertness and performance associated with fatigue, sleep loss, and circadian disruption are issues faced by a diverse range of shiftwork operations. During STS operations, MOD personnel provide 24 hr. coverage of critical tasks. A joint JSC and ARC project was undertaken to examine these issues in flight controllers during MOD shiftwork operations. An initial operational test of procedures and measures was conducted during STS-53 in Dec. 1992. The study measures included a background questionnaire, a subjective daily logbook completed on a 24 hr. basis (to report sleep patterns, work periods, etc.), and an 8 minute performance and mood test battery administered at the beginning, middle, and end of each shift period. Seventeen Flight controllers representing the 3 Orbit shifts participated. The initial results clearly support further data collection during other STS missions to document baseline levels of alertness and performance during MOD shiftwork operations. These issues are especially pertinent for the night shift operations and the acute phase advance required for the transition of day shift personnel into the night for shuttle launch. Implementation and evaluation of the countermeasure strategies to maximize alertness and performance is planned. As STS missions extend to further extended duration orbiters, timelines and planning for 24 circadian disruption will remain highly relevant in the MOD environment.

A methodology for selective removal of orbital debris

NASA Technical Reports Server (NTRS)

Ash, R. L.; Odonoghue, P. J.; Chambers, E. J.; Raney, J. P.

1992-01-01

Earth-orbiting objects, large enough to be tracked, were surveyed for possible systematic debris removal. Based upon the statistical collision studies of others, it was determined that objects in orbits approximately 1000 km above the Earth's surface are at greatest collisional risk. Russian C-1B boosters were identified as the most important target of opportunity for debris removal. Currently, more than 100 in tact boosters are orbiting the Earth with apogees between 950 km and 1050 km. Using data provided by Energia USA, specific information on the C-1B booster, in terms of rendezvous and capture strategies, was discussed.

Intial orbit determination results for Jason-1: towards a 1-cm orbit

NASA Technical Reports Server (NTRS)

Haines, B. J.; Haines, B.; Bertiger, W.; Desai, S.; Kuang, D.; Munson, T.; Reichert, A.; Young, L.; Willis, P.

2002-01-01

The U.S/France Jason-1 oceanographic mission is carrying state-of-the-art radiometric tracking systems (GPS and Doris) to support precise orbit determination (POD) requirements. The performance of the systems is strongly reflected in the early POD results. Results of both internal and external (e.g., satellite laser ranging) comparisons support that the 2.5 cm radial Rh4S requirement is being readily met, and provide reasons for optimism that 1 cm can be achieved. We discuss the POD strategy underlying these orbits, as well as the challenging issues that bear on the understanding and characterization of an orbit solution at the l-cm level. We also describe a system for producing science quality orbits in near real time in order to support emerging applications in operational oceanography.

Individual differences and workload effects on strategy adoption in a dynamic task.

PubMed

Moon, Jungaa; Betts, Shawn; Anderson, John R

2013-09-01

The current study investigated the effects of individual differences and workload on strategy adaptivity in a complex, dynamic task called the Space Fortress game (Donchin, 1989). Participants learned to use a strategy of flying a ship in circles around the fortress in a standard game environment. Once they mastered the strategy, they were assigned to different workload conditions and transferred to a nonstandard environment in which a strong wind was introduced that made it more difficult to achieve a circular orbit. About half of the participants continued with their prior circular strategy while the rest adopted a novel strategy that achieved comparable performance with less effort. With this novel strategy, rather than trying to complete orbits they flew into the wind and then allowed the wind to blow them back to achieve a pendulum-like path. Participants without a working-memory load were more likely to adopt the new strategy. Participants were also more likely to adopt the new strategy if their pattern of behavior exposed them more often to the potential of drifting with the wind. The results indicate that spontaneous changes in strategy occur when people are exposed to the potential of a new strategy and have the cognitive resources to understand its potential. Copyright © 2013 Elsevier B.V. All rights reserved.

The Hubble Space Telescope Servicing Mission 3A Contamination Control Program

NASA Technical Reports Server (NTRS)

Hansen, Patricia A.

2000-01-01

After nearly 10 years on-orbit, the Hubble Space Telescope (HST) external thermal control materials and paint have degraded due to exposure to the low Earth orbit environment. This presented a potentially large on-orbit contamination source (particles and/or debris). Contamination mitigation techniques were developed to augment existing on-orbit servicing contamination controls. They encompassed mission management, crew training, and crew aids and tools. These techniques were successfully employed during the HST Servicing Mission 3A, December 1999.

Pediatric Orbital Fractures

PubMed Central

Oppenheimer, Adam J.; Monson, Laura A.; Buchman, Steven R.

2013-01-01

It is wise to recall the dictum “children are not small adults” when managing pediatric orbital fractures. In a child, the craniofacial skeleton undergoes significant changes in size, shape, and proportion as it grows into maturity. Accordingly, the craniomaxillofacial surgeon must select an appropriate treatment strategy that considers both the nature of the injury and the child's stage of growth. The following review will discuss the management of pediatric orbital fractures, with an emphasis on clinically oriented anatomy and development. PMID:24436730

Control Of Flexible Structures-2 (COFS-2) flight control, structure and gimbal system interaction study

NASA Technical Reports Server (NTRS)

Fay, Stanley; Gates, Stephen; Henderson, Timothy; Sackett, Lester; Kirchwey, Kim; Stoddard, Isaac; Storch, Joel

1988-01-01

The second Control Of Flexible Structures Flight Experiment (COFS-2) includes a long mast as in the first flight experiment, but with the Langley 15-m hoop column antenna attached via a gimbal system to the top of the mast. The mast is to be mounted in the Space Shuttle cargo bay. The servo-driven gimbal system could be used to point the antenna relative to the mast. The dynamic interaction of the Shuttle Orbiter/COFS-2 system with the Orbiter on-orbit Flight Control System (FCS) and the gimbal pointing control system has been studied using analysis and simulation. The Orbiter pointing requirements have been assessed for their impact on allowable free drift time for COFS experiments. Three fixed antenna configurations were investigated. Also simulated was Orbiter attitude control behavior with active vernier jets during antenna slewing. The effect of experiment mast dampers was included. Control system stability and performance and loads on various portions of the COFS-2 structure were investigated. The study indicates possible undesirable interaction between the Orbiter FCS and the flexible, articulated COFS-2 mast/antenna system, even when restricted to vernier reaction jets.

Analyzing costs of space debris mitigation methods

NASA Astrophysics Data System (ADS)

Wiedemann, C.; Krag, H.; Bendisch, J.; Sdunnus, H.

2004-01-01

The steadily increasing number of space objects poses a considerable hazard to all kinds of spacecraft. To reduce the risks to future space missions different debris mitigation measures and spacecraft protection techniques have been investigated during the last years. However, the economic efficiency has not been considered yet in this context. Current studies have the objective to evaluate the mission costs due to space debris in a business as usual (no mitigation) scenario compared to the missions costs considering debris mitigation. The aim is an estimation of the time until the investment in debris mitigation will lead to an effective reduction of mission costs. This paper presents the results of investigations on the key issues of cost estimation for spacecraft and the influence of debris mitigation and shielding on cost. Mitigation strategies like the reduction of orbital lifetime and de- or re-orbit of non-operational satellites are methods to control the space debris environment. These methods result in an increase of costs. In a first step the overall costs of different types of unmanned satellites are analyzed. A selected cost model is simplified and generalized for an application on all operational satellites. In a next step the influence of space debris on cost is treated, if the implementation of mitigation strategies is considered.

ASAP- ARTIFICIAL SATELLITE ANALYSIS PROGRAM

NASA Technical Reports Server (NTRS)

Kwok, J.

1994-01-01

The Artificial Satellite Analysis Program (ASAP) is a general orbit prediction program which incorporates sufficient orbit modeling accuracy for mission design, maneuver analysis, and mission planning. ASAP is suitable for studying planetary orbit missions with spacecraft trajectories of reconnaissance (flyby) and exploratory (mapping) nature. Sample data is included for a geosynchronous station drift cycle study, a Venus radar mapping strategy, a frozen orbit about Mars, and a repeat ground trace orbit. ASAP uses Cowell's method in the numerical integration of the equations of motion. The orbital mechanics calculation contains perturbations due to non-sphericity (up to a 40 X 40 field) of the planet, lunar and solar effects, and drag and solar radiation pressure. An 8th order Runge-Kutta integration scheme with variable step size control is used for efficient propagation. The input includes the classical osculating elements, orbital elements of the sun relative to the planet, reference time and dates, drag coefficient, gravitational constants, and planet radius, rotation rate, etc. The printed output contains Cartesian coordinates, velocity, equinoctial elements, and classical elements for each time step or event step. At each step, selected output is added to a plot file. The ASAP package includes a program for sorting this plot file. LOTUS 1-2-3 is used in the supplied examples to graph the results, but any graphics software package could be used to process the plot file. ASAP is not written to be mission-specific. Instead, it is intended to be used for most planetary orbiting missions. As a consequence, the user has to have some basic understanding of orbital mechanics to provide the correct input and interpret the subsequent output. ASAP is written in FORTRAN 77 for batch execution and has been implemented on an IBM PC compatible computer operating under MS-DOS. The ASAP package requires a math coprocessor and a minimum of 256K RAM. This program was last updated in 1988 with version 2.03. IBM PC is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. Lotus and 1-2-3 are registered trademarks of Lotus Development Corporation.

Electric Propulsion System for Constellation Deployment and Orbit Control of Minisats

NASA Astrophysics Data System (ADS)

Bianco, P.; de Rocco, L.; Lovera, M.

1999-09-01

The late technology developments and the demand for low-cost space missions have raised the interest in small satellites and in their potential use as parts of satellite formations as well as building units of satellite constellations. Formation flying of small satellites can be used to bring in-orbit spares for failed payloads on larger satellites as well as to replace large satellites at all by flying the mission on more small satellites, each carrying a single payload. Small satellites can be used in constellations for scientific missions (e.g. remote sensing) as well as for commercial purposes (e.g. data relay). Yet, "small satellite" doesn't necessarily mean "cheap satellite": cost reduction must be enforced into the space mission design since the very beginning of it, at system level. This usually implies seeking for trade-offs on most expensive system items for a small sat. Among these, we surely have the launch and the onboard propulsion system for orbital manoeuvres and station keeping: the stricter the requirements, the higher the costs. And, when dealing with satellite constellations or formations, orbital requirements can be quite challenging. The system designer is faced with the dilemma on whether to buy a relatively expensive dedicated launch or to have a highly cost-impactive autonomous onboard propulsion system that should perform orbit transfers as well. The present paper, which is an up-to-date version of the one presented at IAF-99, introduces a system based on FEEP (Field Emission Electric Propulsion) technology, featuring low thrust plug-on propulsion units. Thanks to the self-contained concept of FEEP thrusters and to the plug-on feature of the whole system, a very low cost-impactive onboard propulsion system can be implemented in order to serve for both orbital manoeuvres (constellation / formation deployment, orbit rising) and orbit maintenance (drag compensation, station keeping relative to other satellites). Most convenient strategies to operate such propulsion systems with respect to orbital requirements, principal design drivers and sizing methods are presented and discussed as well as practical up-to-date case study results performed at Carlo Gavazzi Space.

Seasat-A attitude control system

NASA Technical Reports Server (NTRS)

Weiss, R.; Rodden, J. J.; Hendricks, R. J.

1977-01-01

The Seasat-A attitude control system controls the attitude of the satellite system during injection into final circular orbit after Atlas boost, during orbit adjust and trim phases, and throughout the 3-year mission. Ascent and injection guidance and attitude control are provided by the Agena spacecraft with a gyrocompassed mass expulsion system. On-orbit attitude control functions are performed by a system that has its functional roots in the gravity-gradient momentum bias technology. The paper discusses hardware, control laws, and simulation results.

Chandra X-Ray Observatory Pointing Control System Performance During Transfer Orbit and Initial On-Orbit Operations

NASA Technical Reports Server (NTRS)

Quast, Peter; Tung, Frank; West, Mark; Wider, John

2000-01-01

The Chandra X-ray Observatory (CXO, formerly AXAF) is the third of the four NASA great observatories. It was launched from Kennedy Space Flight Center on 23 July 1999 aboard the Space Shuttle Columbia and was successfully inserted in a 330 x 72,000 km orbit by the Inertial Upper Stage (IUS). Through a series of five Integral Propulsion System burns, CXO was placed in a 10,000 x 139,000 km orbit. After initial on-orbit checkout, Chandra's first light images were unveiled to the public on 26 August, 1999. The CXO Pointing Control and Aspect Determination (PCAD) subsystem is designed to perform attitude control and determination functions in support of transfer orbit operations and on-orbit science mission. After a brief description of the PCAD subsystem, the paper highlights the PCAD activities during the transfer orbit and initial on-orbit operations. These activities include: CXO/IUS separation, attitude and gyro bias estimation with earth sensor and sun sensor, attitude control and disturbance torque estimation for delta-v burns, momentum build-up due to gravity gradient and solar pressure, momentum unloading with thrusters, attitude initialization with star measurements, gyro alignment calibration, maneuvering and transition to normal pointing, and PCAD pointing and stability performance.

Deploying the NASA Meter Class Autonomous Telescope (MCAT) on Ascension Island

NASA Technical Reports Server (NTRS)

Lederer, S. M.; Pace, L.; Hickson, P.; Cowardin, H. M.; Frith, J.; Buckalew, B.; Glesne, T.; Maeda, R.; Douglas, D.; Nishimoto, D.

2015-01-01

NASA has successfully constructed the 1.3m Meter Class Autonomous Telescope (MCAT) facility on Ascension Island in the South Atlantic Ocean. MCAT is an optical telescope designed specifically to collect ground-based data for the statistical characterization of orbital debris ranging from Low Earth Orbit (LEO) through Middle Earth Orbits (MEO) and beyond to Geo Transfer and Geosynchronous Orbits (GTO/GEO). The location of Ascension Island has two distinct advantages. First, the near-equatorial location fills a significant longitudinal gap in the Ground-based Electro-Optical Deep Space Surveillance (GEODSS) network of telescopes, and second, it allows access to objects in Low Inclination Low-Earth Orbits (LILO). The MCAT facility will be controlled by a sophisticated software suite that operates the dome and telescope, assesses sky and weather conditions, conducts all necessary calibrations, defines an observing strategy (as dictated by weather, sky conditions and the observing plan for the night), and carries out the observations. It then reduces the collected data via four primary observing modes ranging from tracking previously cataloged objects to conducting general surveys for detecting uncorrelated debris. Nightly observing plans, as well as the resulting text file of reduced data, will be transferred to and from Ascension, respectively, via a satellite connection. Post-processing occurs at NASA Johnson Space Center. Construction began in September, 2014 with dome and telescope installation occurring in April through early June, 2015. First light was achieved in June, 2015. Acceptance testing, full commissioning, and calibration of this soon-to-be fully autonomous system commenced in summer 2015. The initial characterization of the system from these tests is presented herein.

Deploying the NASA Meter Class Autonomous Telescope (MCAT) on Ascension Island

NASA Astrophysics Data System (ADS)

Lederer, S.; Pace, L. F.; Hickson, P.; Glesne, T.; Cowardin, H. M.; Frith, J. M.; Buckalew, B.; Maeda, R.; Douglas, D.; Nishimoto, D.

NASA has successfully constructed the 1.3m Meter Class Autonomous Telescope (MCAT) facility on Ascension Island in the South Atlantic Ocean. MCAT is an optical telescope designed specifically to collect ground-based data for the statistical characterization of orbital debris ranging from Low Earth Orbit (LEO) through Middle Earth Orbits (MEO) and beyond to Geo Transfer and Geosynchronous Orbits (GTO/GEO). The location of Ascension Island has two distinct advantages. First, the near-equatorial location fills a significant longitudinal gap in the Ground-based Electro-Optical Deep Space Surveillance (GEODSS) network of telescopes, and second, it allows access to objects in Low Inclination Low-Earth Orbits (LILO). The MCAT facility will be controlled by a sophisticated software suite that operates the dome and telescope, assesses sky and weather conditions, conducts all necessary calibrations, defines an observing strategy (as dictated by weather, sky conditions, and the observing plan for the night), and carries out the observations. It then reduces the collected data via four primary observing modes ranging from tracking previously cataloged objects to conducting general surveys for detecting uncorrelated debris. Nightly observing plans, as well as the resulting text file of reduced data, will be transferred to and from Ascension, respectively, via a satellite connection. Post-processing occurs at NASA Johnson Space Center. Construction began in September, 2014 with dome and telescope installation occurring in April through early June, 2015. First light was achieved in June, 2015. Acceptance testing, full commissioning, and calibration of this soon-to-be fully autonomous system commenced in summer 2015. The initial characterization of the system from these tests is presented herein.

Designing capture trajectories to unstable periodic orbits around Europa

NASA Technical Reports Server (NTRS)

Russell, Ryan P.; Lam, Try

2006-01-01

The hostile environment of third body perturbations restricts a mission designer's ability to find well-behaved reproducible capture trajectories when dealing with limited control authority as is typical with low-thrust missions. The approach outlined in this paper confronts this shortcoming by utilizing dynamical systems theory and an extensive preexisting database of Restricted Three Body Problem (RTBP) periodic orbits. The stable manifolds of unstable periodic orbits are utilized to attract a spacecraft towards Europa. By selecting an appropriate periodic orbit, a mission designer can control important characteristics of the captured state including stability, minimum altitudes, characteristic inclinations, and characteristic radii among others. Several free parameters are optimized in the non-trivial mapping from the RTBP to a more realistic model. Although the ephemeris capture orbit is ballistic by design, low-thrust is used to target the state that leads to the capture orbit, control the spacecraft after arriving on the unstable quasi-periodic orbit, and begin the spiral down towards the science orbit. The approach allows a mission designer to directly target fuel efficient captures at Europa in an ephemeris model. Furthermore, it provides structure and controllability to the design of capture trajectories that reside in a chaotic environment.

The Hot Orbit: Orbital Cellulitis

PubMed Central

Chaudhry, Imtiaz A.; Al-Rashed, Waleed; Arat, Yonca O.

2012-01-01

Orbital cellulitis is an uncommon condition previously associated with severe complications. If untreated, orbital cellulitis can be potentially sight and life threatening. It can affect both adults and children but has a greater tendency to occur in the pediatric age group. The infection most commonly originates from sinuses, eyelids or face, retained foreign bodies, or distant soources by hematogenous spread. It is characterized by eyelid edema, erythema, chemosis, proptosis, blurred vision, fever, headache, and double vision. A history of upper respiratory tract infection prior to the onset is very common especially in children. In the era prior to antibiotics, vision loss from orbital cellulitis was a dreaded complication. Currently, imaging studies for detection of orbital abcess, the use of antibiotics and early drainage have mitigated visual morbidity significantly. The purpose of this review is to describe current investigative strategies and management options in the treatment of orbital cellulitis, establish their effectiveness and possible complications due to late intervention. PMID:22346113

Orbiting Geophysical Observatory Attitude Control Subsystem Design Survey. NASA/ERC Design Criteria Program, Guidance and Control

NASA Technical Reports Server (NTRS)

Mc Kenna, K. J.; Schmeichel, H.

1968-01-01

This design survey summarizes the history of the Orbiting Geophysical Observatories' (OGO) Attitude Control Subsystem (ACS) from the proposal phase through current flight experience. Problems encountered in design, fabrication, test, and on orbit are discussed. It is hoped that the experiences of the OGO program related here will aid future designers.

Hypersonic vehicle control law development using H(infinity) and micron-synthesis

NASA Technical Reports Server (NTRS)

Gregory, Irene M.; Mcminn, John D.; Shaughnessy, John D.; Chowdhry, Rajiv S.

1993-01-01

Hypersonic vehicle control law development using H(infinity) and mu-synthesis is discussed. Airbreathing SSTO vehicles has a mutli-faceted mission that includes orbital operations, as well as re-entry and descent culminating in horizontal landing. However, the most challenging part of the operations is the ascent to orbit. The airbreathing propulsion requires lengthy atmospheric flight that may last as long as 30 minutes and take the vehicle half way around the globe. The vehicles's ascent is characterized by tight payload to orbit margins which translate into minimum fuel orbit as the performance criteria. Issues discussed include: SSTO airbreathing vehicle issues; control system performance requirements; robust control law framework; H(infinity) controller frequency analysis; and mu controller frequency analysis.

Fuel-efficient feedback control of orbital motion around irregular-shaped asteroids

NASA Astrophysics Data System (ADS)

Winkler, Timothy Michael

Unmanned probes are the primary technologies used when exploring celestial bodies in our solar system. As these deep space exploration missions are becoming more and more complex, there is a need for advanced autonomous operation capabilities in order to meet mission objectives. These autonomous capabilities are required as ground-based guidance and navigation commands will not be able to be issued in real time due to the large distance from the Earth. For long-duration asteroid exploration missions, this also entails how to keep the spacecraft around or on the body in order for the mission to be successfully completed. Unlike with larger bodies such as planets, though, the dynamical environment around these smaller bodies can be difficult to characterize. The weak gravitational fields are not uniform due to irregular shapes and non-homogeneous mass distribution, especially when orbiting in close-proximity to the body. On top of that, small perturbation forces such as solar radiation pressure can be strong enough to destabilize an orbit around an asteroid. The best solution for keeping a spacecraft in orbit about a small body is to implement some form of control technique. With conventional propulsion thrusters, active control algorithms tend to have a higher than acceptable propellant requirements for long-duration asteroid exploration missions, which has led to much research being devoted to finding open-loop solutions to long-term stable orbits about small bodies. These solutions can prove to be highly sensitive to the orbit's initial conditions, making them potentially unreliable in the presence of orbit injection errors. This research investigates a fuel-efficient, active control scheme to safely control a spacecraft's orbit in close-proximity to an asteroid. First, three different gravitational models capable of simulating the non-homogeneous gravity fields of asteroids are presented: the polyhedron gravity shape model, a spherical harmonics expansion, and an inertia dyadic gravity model. Then a simple feedback controller augmented by a disturbance-accommodating filter is employed to ensure orbital stability. Using these models and controller, several orbiting cases as well as body-frame hovering are investigated to test the viability and fuel-efficiency of the proposed control system. The ultimate goal is to design an active orbit control system with minimum DeltaV expenditure.

Extracting More Information from Passive Optical Tracking Observations for Reliable Orbit Element Generation

NASA Astrophysics Data System (ADS)

Bennett, J.; Gehly, S.

2016-09-01

This paper presents results from a preliminary method for extracting more orbital information from low rate passive optical tracking data. An improvement in the accuracy of the observation data yields more accurate and reliable orbital elements. A comparison between the orbit propagations from the orbital element generated using the new data processing method is compared with the one generated from the raw observation data for several objects. Optical tracking data collected by EOS Space Systems, located on Mount Stromlo, Australia, is fitted to provide a new orbital element. The element accuracy is determined from a comparison between the predicted orbit and subsequent tracking data or reference orbit if available. The new method is shown to result in a better orbit prediction which has important implications in conjunction assessments and the Space Environment Research Centre space object catalogue. The focus is on obtaining reliable orbital solutions from sparse data. This work forms part of the collaborative effort of the Space Environment Management Cooperative Research Centre which is developing new technologies and strategies to preserve the space environment (www.serc.org.au).

A MATLAB based Distributed Real-time Simulation of Lander-Orbiter-Earth Communication for Lunar Missions

NASA Astrophysics Data System (ADS)

Choudhury, Diptyajit; Angeloski, Aleksandar; Ziah, Haseeb; Buchholz, Hilmar; Landsman, Andre; Gupta, Amitava; Mitra, Tiyasa

Lunar explorations often involve use of a lunar lander , a rover [1],[2] and an orbiter which rotates around the moon with a fixed radius. The orbiters are usually lunar satellites orbiting along a polar orbit to ensure visibility with respect to the rover and the Earth Station although with varying latency. Communication in such deep space missions is usually done using a specialized protocol like Proximity-1[3]. MATLAB simulation of Proximity-1 have been attempted by some contemporary researchers[4] to simulate all features like transmission control, delay etc. In this paper it is attempted to simulate, in real time, the communication between a tracking station on earth (earth station), a lunar orbiter and a lunar rover using concepts of Distributed Real-time Simulation(DRTS).The objective of the simulation is to simulate, in real-time, the time varying communication delays associated with the communicating elements with a facility to integrate specific simulation modules to study different aspects e.g. response due to a specific control command from the earth station to be executed by the rover. The hardware platform comprises four single board computers operating as stand-alone real time systems (developed by MATLAB xPC target and inter-networked using UDP-IP protocol). A time triggered DRTS approach is adopted. The earth station, the orbiter and the rover are programmed as three standalone real-time processes representing the communicating elements in the system. Communication from one communicating element to another constitutes an event which passes a state message from one element to another, augmenting the state of the latter. These events are handled by an event scheduler which is the fourth real-time process. The event scheduler simulates the delay in space communication taking into consideration the distance between the communicating elements. A unique time synchronization algorithm is developed which takes into account the large latencies in space communication. The DRTS setup thus developed serves as an important and inexpensive test bench for trying out remote controlled applications on the rover, for example, from an earth station. The simulation is modular and the system is composable. Each of the processes can be aug-mented with relevant simulation modules that handle the events to simulate specific function-alities. With stringent energy saving requirements on most rovers, such a simulation set up, for example, can be used to design optimal rover movement control strategies from the orbiter in conjunction with autonomous systems on the rover itself. References 1. Lunar and Planetary Department, Moscow University, Lunokhod 1, "http://selena.sai.msu.ru/Home/Spa 2. NASA History Office, Guidelines for Advanced Manned Space Vehicle Program, "http://history.nasa.gov 35ann/AMSVPguidelines/top.htm" 3. Consultative Committee For Space Data Systems, "Proximity-1 Space Link Protocol" CCSDS 211.0-B-1 Blue Book. October 2002. 4. Segui, J. and Jennings, E., "Delay Tolerant Networking-Bundle Protocol Simulation", in Proceedings of the 2nd IEEE International Conference on Space Mission Challenges for Infor-mation Technology, 2006.

General view of the aft section of the Orbiter Discovery ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view of the aft section of the Orbiter Discovery in the Vehicle Assembly Building at NASA's Kennedy Space Center. Note the main engines and Orbiter Maneuvering System/Reaction Control System pods are removed in this photo. The flexible hoses protruding from the starboard aft section are to control temperature, humidity and pressure in the orbiter's void spaces during its down time. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

General view of the aft, starboard section of the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view of the aft, starboard section of the Orbiter Discovery in the Vehicle Assembly Building at NASA's Kennedy Space Center. Note the main engines and Orbiter Maneuvering System/Reaction Control System pods are removed in this photo. The flexible hoses protruding from the starboard aft section are to control temperature, humidity and pressure in the orbiter's void spaces during its down time. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Stationkeeping Approach for the Microwave Anisotropy Probe (MAP)

NASA Technical Reports Server (NTRS)

Rohrbaugh, Dave; Schiff, Conrad

2002-01-01

The Microwave Anisotropy Probe was successfully launched on June 30, 2001 and placed into a Lissajous orbit about the L2 Sun-Earth-Moon libration point. However, the L2 libration point is unstable which necessitates occasional stationkeeping maneuvers in order to maintain the spacecraft s Lissajous orbit. Analyses were performed in order to develop a feasible L2 stationkeeping strategy for the MAP mission. The resulting strategy meets the allotted fuel budget, allowing for enough fuel to handle additional he1 taxes, while meeting the attitude requirements for the maneuvers. Results from the first two stationkeeping maneuvers are included.

A Preliminary Data Model for Orbital Flight Dynamics in Shuttle Mission Control

NASA Technical Reports Server (NTRS)

ONeill, John; Shalin, Valerie L.

2000-01-01

The Orbital Flight Dynamics group in Shuttle Mission Control is investigating new user interfaces in a project called RIOTS [RIOTS 2000]. Traditionally, the individual functions of hardware and software guide the design of displays, which results in an aggregated, if not integrated interface. The human work system has then been designed and trained to navigate, operate and integrate the processors and displays. The aim of RIOTS is to reduce the cognitive demands of the flight controllers by redesigning the user interface to support the work of the flight controller. This document supports the RIOTS project by defining a preliminary data model for Orbital Flight Dynamics. Section 2 defines an information-centric perspective. An information-centric approach aims to reduce the cognitive workload of the flight controllers by reducing the need for manual integration of information across processors and displays. Section 3 describes the Orbital Flight Dynamics domain. Section 4 defines the preliminary data model for Orbital Flight Dynamics. Section 5 examines the implications of mapping the data model to Orbital Flight Dynamics current information systems. Two recurring patterns are identified in the Orbital Flight Dynamics work the iteration/rework cycle and the decision-making/information integration/mirroring role relationship. Section 6 identifies new requirements on Orbital Flight Dynamics work and makes recommendations based on changing the information environment, changing the implementation of the data model, and changing the two recurring patterns.

Enhanced orbit determination filter sensitivity analysis: Error budget development

NASA Technical Reports Server (NTRS)

Estefan, J. A.; Burkhart, P. D.

1994-01-01

An error budget analysis is presented which quantifies the effects of different error sources in the orbit determination process when the enhanced orbit determination filter, recently developed, is used to reduce radio metric data. The enhanced filter strategy differs from more traditional filtering methods in that nearly all of the principal ground system calibration errors affecting the data are represented as filter parameters. Error budget computations were performed for a Mars Observer interplanetary cruise scenario for cases in which only X-band (8.4-GHz) Doppler data were used to determine the spacecraft's orbit, X-band ranging data were used exclusively, and a combined set in which the ranging data were used in addition to the Doppler data. In all three cases, the filter model was assumed to be a correct representation of the physical world. Random nongravitational accelerations were found to be the largest source of error contributing to the individual error budgets. Other significant contributors, depending on the data strategy used, were solar-radiation pressure coefficient uncertainty, random earth-orientation calibration errors, and Deep Space Network (DSN) station location uncertainty.

Space Technology 5: Changing the Mission Design without Changing the Hardware

NASA Technical Reports Server (NTRS)

Carlisle, Candace C.; Webb, Evan H.; Slavin, James A.

2005-01-01

The Space Technology 5 (ST-5) Project is part of NASA's New Millennium Program. The validation objectives are to demonstrate the research-quality science capability of the ST-5 spacecraft; to operate the three spacecraft as a constellation; and to design, develop, test and flight-validate three capable micro-satellites with new technologies. A three-month flight demonstration phase is planned, beginning in March 2006. This year, the mission was re-planned for a Pegasus XL dedicated launch into an elliptical polar orbit (instead of the Originally-planned Geosynchronous Transfer Orbit.) The re-plan allows the mission to achieve the same high-level technology validation objectives with a different launch vehicle. The new mission design involves a revised science validation strategy, a new orbit and different communication strategy, while minimizing changes to the ST-5 spacecraft itself. The constellation operations concepts have also been refined. While the system engineers, orbit analysts, and operations teams were re-planning the mission, the implementation team continued to make progress on the flight hardware. Most components have been delivered, and the first spacecraft is well into integration and test.

Orbital transfer and release of tethered payloads. Continuation of investigation of electrodynamic stabilization and control of long orbiting tethers Martinez-Sanchez, Manuel

NASA Technical Reports Server (NTRS)

Colombo, G.; Grossi, M. D.; Arnold, D.

1983-01-01

The effect of reeling operations on the orbital altitude of the tether system and the development of control laws to minimize tether rebound upon payload release were studied. The use of the tether for LEO/GEO payload orbital transfer was also investigated. It was concluded that (1) reeling operations can contribute a significant amount of energy to the orbit of the system and should be considered in orbit calculations and predictions, (2) deployment of payloads, even very large payloads, using tethers is a practical and fully stable operation, (3) tether augmented LEO/GEO transfer operations yield useful payload gains under the practical constraint of fixed size OTV's, and (4) orbit to orbit satellite retrieval is limited by useful revisit times to orbital inclinations of less than forty-five degrees.

Time-free transfers between libration-point orbits in the elliptic restricted problem

NASA Astrophysics Data System (ADS)

Howell, K. C.; Hiday-Johnston, L. A.

This work is part of a larger research effort directed toward the formulation of a strategy to design optimal time-free impulsive transfers between three-dimensional libration-point orbits in the vicinity of the interior LI libration point of the Sun-Earth/Moon barycenter system. Inferior transfers that move a spacecraft from a large halo orbit to a smaller halo orbit are considered here. Primer vector theory is applied to non-optimal impulsive trajectories in the elliptic restricted three-body problem in order to establish whether the implementation of a coast in the initial orbit, a coast in the final orbit, or dual coasts accomplishes a reduction in fuel expenditure. The addition of interior impulses is also considered. Results indicate that a substantial savings in fuel can be achieved by the allowance for coastal periods on the specified libration-point orbits. The resulting time-free inferior transfers are compared to time-free superior transfers between halo orbits of equal z-amplitude separation.

Time-free transfers between libration-point orbits in the elliptic restricted problem

NASA Astrophysics Data System (ADS)

Howell, K. C.; Hiday, L. A.

1992-08-01

This work is directed toward the formulation of a strategy to design optimal time-free impulsive transfers between 3D libration-point orbits in the vicinity of the interior L1 libration point of the sun-earth/moon barycenter system. Inferior transfers that move a spacecraft from a large halo orbit to a smaller halo orbit are considered here. Primer vector theory is applied to nonoptimal impulsive trajectories in the elliptic restricted three-body problem in order to establish whether the implementation of a coast in the initial orbit, a coast in the final orbit, or dual coasts accomplishes a reduction in fuel expenditure. The addition of interior impulses is also considered. Results indicate that a substantial savings in fuel can be achieved by the allowance for coastal periods on the specified libration-point orbits. The resulting time-free inferior transfers are compared to time-free superior transfers between halo orbits of equal z-amplitude separation.

Lunar prospector mission design and trajectory support

NASA Technical Reports Server (NTRS)

Lozier, David; Galal, Ken; Folta, David; Beckman, Mark

1998-01-01

The Lunar Prospector mission is the first dedicated NASA lunar mapping mission since the Apollo Orbiter program which was flown over 25 years ago. Competitively selected under the NASA Discovery Program, Lunar Prospector was launched on January 7, 1998 on the new Lockheed Martin Athena 2 launch vehicle. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with three scheduled orbit correction maneuvers to remove launch and cislunar injection errors prior to lunar insertion. At lunar encounter, a series of three lunar orbit insertion maneuvers and a small circularization burn were executed to achieve a 100 km altitude polar mapping orbit. This paper will present the design of the Lunar Prospector transfer, lunar insertion and mapping orbits, including maneuver and orbit determination strategies in the context of mission goals and constraints. Contingency plans for handling transfer orbit injection and lunar orbit insertion anomalies are also summarized. Actual flight operations results are discussed and compared to pre-launch support analysis.

Space Operations Center, shuttle interaction study, volume 1

NASA Technical Reports Server (NTRS)

1981-01-01

The feasibility of the shuttle remote manipulator system (SRMS)-aided space operations center (SOC)/orbiter berthing was evaluated to determine: (1) whether the initial rates between the SOC and the orbiter can be removed by the arm; (2) what is the best strategy to be used; (3) whether the post-capture and maneuvering loads are within the capability of the SRMS; (4) can the SOC berthing port be brought in the immediate proximity of the orbiter berthing port; and (5) what is the best way to remove the residual relative motions. Various notational conventions are established and various important locations on the orbiter and SOC structures are defined. Reference frames are defined together with the mass properties of both the SOC and the orbiter.

Satellite tour design for the Cassini mission. [Saturn

NASA Technical Reports Server (NTRS)

Diehl, R. E.

1985-01-01

The mission design for the Cassini satellite tour of Saturn is described. The use of Titan encounters for gravity-assisted trajectory shaping to satisfy atmosphere, magnetosphere, and satellite science requirements is discussed. Three candidate satellite tours that illustrate the tradeoffs between different strategies to satisfy the science objectives are presented. Tour characteristics for a 4yr mission include at least 30 Titan encounters to provide a minimum of 90 deg orbit rotation for either a magnetotail petal orbit or noon petal orbit; evenly distributed Titan ground tracks for radar mapping coverage; at least 2 Iapetus and 1 Enceladus close flybys; inclined orbits during the first 3 yr to provide Earth and Sun occultations by Saturn and its rings; and Polar orbits during the fourth year.

Earth-Moon Libration Point Orbit Stationkeeping: Theory, Modeling and Operations

NASA Technical Reports Server (NTRS)

Folta, David C.; Pavlak, Thomas A.; Haapala, Amanda F.; Howell, Kathleen C.; Woodard, Mark A.

2013-01-01

Collinear Earth-Moon libration points have emerged as locations with immediate applications. These libration point orbits are inherently unstable and must be maintained regularly which constrains operations and maneuver locations. Stationkeeping is challenging due to relatively short time scales for divergence effects of large orbital eccentricity of the secondary body, and third-body perturbations. Using the Acceleration Reconnection and Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) mission orbit as a platform, the fundamental behavior of the trajectories is explored using Poincare maps in the circular restricted three-body problem. Operational stationkeeping results obtained using the Optimal Continuation Strategy are presented and compared to orbit stability information generated from mode analysis based in dynamical systems theory.

The research of the coupled orbital-attitude controlled motion of celestial body in the neighborhood of the collinear libration point L1

NASA Astrophysics Data System (ADS)

Shmyrov, A.; Shmyrov, V.; Shymanchuk, D.

2017-10-01

This article considers the motion of a celestial body within the restricted three-body problem of the Sun-Earth system. The equations of controlled coupled attitude-orbit motion in the neighborhood of collinear libration point L1 are investigated. The translational orbital motion of a celestial body is described using Hill's equations of circular restricted three-body problem of the Sun-Earth system. Rotational orbital motion is described using Euler's dynamic equations and quaternion kinematic equation. We investigate the problem of stability of celestial body rotational orbital motion in relative equilibrium positions and stabilization of celestial body rotational orbital motion with proposed control laws in the neighborhood of collinear libration point L1. To study stabilization problem, Lyapunov function is constructed in the form of the sum of the kinetic energy and special "kinematic function" of the Rodriguez-Hamiltonian parameters. Numerical modeling of the controlled rotational motion of a celestial body at libration point L1 is carried out. The numerical characteristics of the control parameters and rotational motion are given.

Independent Orbiter Assessment (IOA): Analysis of the reaction control system, volume 1

NASA Technical Reports Server (NTRS)

Burkemper, V. J.; Haufler, W. A.; Odonnell, R. A.; Paul, D. J.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results for the Reaction Control System (RCS). The purpose of the RCS is to provide thrust in and about the X, Y, Z axes for External Tank (ET) separation; orbit insertion maneuvers; orbit translation maneuvers; on-orbit attitude control; rendezvous; proximity operations (payload deploy and capture); deorbit maneuvers; and abort attitude control. The RCS is situated in three independent modules, one forward in the orbiter nose and one in each OMS/RCS pod. Each RCS module consists of the following subsystems: Helium Pressurization Subsystem; Propellant Storage and Distribution Subsystem; Thruster Subsystem; and Electrical Power Distribution and Control Subsystem. Of the failure modes analyzed, 307 could potentially result in a loss of life and/or loss of vehicle.

RCS jet-flow field interaction effects on the aerodynamics of the space shuttle orbiter

NASA Technical Reports Server (NTRS)

Rausch, J. R.; Roberge, A. M.

1973-01-01

A study was conducted to determine the external effects caused by operation of the reaction control system during entry of the space shuttle orbiter. The effects of jet plume-external flow interactions were emphasized. Force data were obtained for the basic airframe characteristics plus induced effects when the reaction control system is operating. Resulting control amplification and/or coupling were derived and their effects on the aerodynamic stability and control of the orbiter and the reaction control system thrust were determined.

Strategies for the return of science data from in situ vehicles at Titan

NASA Astrophysics Data System (ADS)

Spilker, T. R.; Reh, K. R.; Erd, C.; Elliott, J. O.; Mohr, D.; Strange, N. J.

2009-04-01

Collaborative studies of the Titan Saturn System Mission (TSSM) in 2008 by ESA and NASA have included examination of strategies for optimizing the science return from that mission concept's proposed in situ elements. The current baselined mission concept calls for an orbiter provided and launched by NASA that would deliver to Titan and support two ESA-provided in situ elements, a lake lander whose science mission duration would be about nine hours, and a montgolfière (hot-air balloon) that would operate at ~10 km altitude in Titan's lower atmosphere for 6-12 months. This architecture has much in common with the highly successful Cassini-Huygens mission. The short-lived lake lander in particular would have a mission profile very similar to that of the Huygens probe, with all science data communications occurring while the NASA orbiter is relatively near Titan. Practical mission profile options for the montgolfière include extended periods when the NASA orbiter is farther from Titan, reducing data rates. Over long periods of time the montgolfière cannot be considered fixed over one location on Titan's surface, and in fact is expected to circumnavigate Titan in less than six months. Thus the schedule of communications windows between the in situ elements and the orbiter cannot be precisely determined far in advance, varying as the balloon literally "rides the wind". Other issues played critical roles in evaluating the many options available early in the studies. Some options for the timing of delivery of the in situ elements yielded more mass capability available for those elements, but their reduced data return due to orbit geometry outweighs the added mass capability. Another delivery option, delivery from Titan orbit, yields reduced delivery mass capability but was thought (before studies) to offer better data relay capability. Studies revealed that this strategy actually decreases the return from the lake lander as compared to options delivering the in situ elements from hyperbolic flybys. This presentation will describe options examined in the TSSM communications strategy studies. Particular attention is given to that chosen for the baseline strategy, with potential returned data volumes that provide generous margins over anticipated data requirements. Many of the results are not unique to Titan alone, but are applicable to in situ missions at any satellite of a giant planet. These collaborative studies were funded by, and performed under the cognizance of, NASA and ESA.

Risk Analysis of On-Orbit Spacecraft Refueling Concepts

NASA Technical Reports Server (NTRS)

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

2010-01-01

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

Proceedings of the 20th International Symposium on Space Flight Dynamics

NASA Technical Reports Server (NTRS)

Woodard, Mark (Editor); Stengle, Tom (Editor)

2007-01-01

Topics include: Measuring Image Navigation and Registration Performance at the 3-Sigma Level Using Platinum Quality Landmarks; Flight Dynamics Performances of the MetOp A Satellite during the First Months of Operations; Visual Navigation - SARE Mission; Determining a Method of Enabling and Disabling the Integral Torque in the SDO Science and Inertial Mode Controllers; Guaranteeing Pointing Performance of the SDO Sun-Pointing Controllers in Light of Nonlinear Effects; SDO Delta H Mode Design and Analysis; Observing Mode Attitude Controller for the Lunar Reconnaissance Orbiter; Broken-Plane Maneuver Applications for Earth to Mars Trajectories; ExoMars Mission Analysis and Design - Launch, Cruise and Arrival Analyses; Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance; Mars Reconnaissance Orbiter Interplanetary Cruise Navigation; Motion Parameters Determination of the SC and Phobos in the Project Phobos-Grunt; GRAS NRT Precise Orbit Determination: Operational Experience; Orbit Determination of LEO Satellites for a Single Pass through a Radar: Comparison of Methods; Orbit Determination System for Low Earth Orbit Satellites; Precise Orbit Determination for ALOS; Anti-Collision Function Design and Performances of the CNES Formation Flying Experiment on the PRISMA Mission; CNES Approaching Guidance Experiment within FFIORD; Maneuver Recovery Analysis for the Magnetospheric Multiscale Mission; SIMBOL-X: A Formation Flying Mission on HEO for Exploring the Universe; Spaceborne Autonomous and Ground Based Relative Orbit Control for the TerraSAR-X/TanDEM-X Formation; First In-Orbit Experience of TerraSAR-X Flight Dynamics Operations; Automated Target Planning for FUSE Using the SOVA Algorithm; Space Technology 5 Post-Launch Ground Attitude Estimation Experience; Standardizing Navigation Data: A Status Update; and A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer.

Formation Flying With Decentralized Control in Libration Point Orbits

NASA Technical Reports Server (NTRS)

Folta, David; Carpenter, J. Russell; Wagner, Christoph

2000-01-01

A decentralized control framework is investigated for applicability of formation flying control in libration orbits. The decentralized approach, being non-hierarchical, processes only direct measurement data, in parallel with the other spacecraft. Control is accomplished via linearization about a reference libration orbit with standard control using a Linear Quadratic Regulator (LQR) or the GSFC control algorithm. Both are linearized about the current state estimate as with the extended Kalman filter. Based on this preliminary work, the decentralized approach appears to be feasible for upcoming libration missions using distributed spacecraft.

Drag De-Orbit Device (D3): A Retractable Device for CubeSat Attitude and Orbit Control using Aerodynamic Forces

NASA Technical Reports Server (NTRS)

Guglielmo, David; Omar, Sanny R.; Bevilacqua, Riccardo

2017-01-01

The increasing number of CubeSats being launched has raised concerns about orbital debris since most of these satellites have no means of active orbit control. Some technologies exist to increase the surface area of a CubeSat and expedite de-orbit due to aerodynamic drag in low Earth orbit, but most of these devices cannot be retracted and hence cannot be used for orbital maneuvering. This paper discusses the De-Orbit Drag Device (D3) module that is capable of de-orbiting a 12U, 15kg CubeSat from a 700 km circular orbit in under 25 years and can be deployed and retracted to modulate the aerodynamic drag force experienced by the satellite. This facilitates orbital maneuvering using aerodynamic drag and the active targeting of a de-orbit location. In addition, the geometry of this drag device provides 3-axis attitude stabilization of the host CubeSat using aerodynamic and gravity gradient torques which is useful for many missions and provides a predictable aerodynamic profile for use in orbital maneuvering algorithms.

ERTS-1 flight evaluation report, 23 April - 23 July 1973

NASA Technical Reports Server (NTRS)

1973-01-01

The flight performance of the ERTS-1 satellite is analyzed for orbits 3810 to 5100. Systems analyzed include: orbital parameters, power subsystem, attitude control, telemetry, orbit adjust, thermal control, and data collection. Documents and reports related to the evaluation are also included.

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

NASA Technical Reports Server (NTRS)

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.

Transition from orbital liquid to Jahn-Teller insulator in orthorhombic perovskites RTiO3.

PubMed

Cheng, J-G; Sui, Y; Zhou, J-S; Goodenough, J B; Su, W H

2008-08-22

Following the same strategy used for RVO3, thermal conductivity measurements have been made on a series of single-crystal perovskites RTiO3 (R=La,Nd,...,Yb). Results reveal explicitly a transition from an orbital liquid to an orbitally ordered phase at a magnetic transition temperature, which is common for both the antiferromagnetic and ferromagnetic phases in the phase diagram of RTiO3. This spin/orbital transition is consistent with the mode softening at T_{N} in antiferromagnetic LaTiO3 and is supported by an anomalous critical behavior at T_{c} in ferromagnetic YTiO3.

Removing the solar exclusion with high altitude satellites [Orbital strategies to mitigate the Solar Exclusion Effect on Space-Based Observation of the Geosynchronous Belt

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

Vallado, David A.; Cefola, Paul J.; Kiziah, Rex R.

Here, observing geosynchronous satellites has numerous applications. Lighting conditions near the equinoxes routinely cause problems for traditional observations of sensors near the equator – the solar exclusion. We investigate using sensors on satellites (in polar and high- altitude orbits) to observe satellites that are in geosynchronous orbit. It is hoped that these satellite configurations will alleviate many of these problems. Assessing the orbit insertion and station-keeping requirements are important to understand. We summarize the literature to understand the relevant perturbing forces and assess the delta-v requirements.

Removing the solar exclusion with high altitude satellites [Orbital strategies to mitigate the Solar Exclusion Effect on Space-Based Observation of the Geosynchronous Belt

DOE PAGES

Vallado, David A.; Cefola, Paul J.; Kiziah, Rex R.; ...

2016-09-09

Here, observing geosynchronous satellites has numerous applications. Lighting conditions near the equinoxes routinely cause problems for traditional observations of sensors near the equator – the solar exclusion. We investigate using sensors on satellites (in polar and high- altitude orbits) to observe satellites that are in geosynchronous orbit. It is hoped that these satellite configurations will alleviate many of these problems. Assessing the orbit insertion and station-keeping requirements are important to understand. We summarize the literature to understand the relevant perturbing forces and assess the delta-v requirements.

Collisional cascading - The limits of population growth in low earth orbit

NASA Technical Reports Server (NTRS)

Kessler, Donald J.

1991-01-01

Random collisions between made-made objects in earth orbit will lead to a significant source of orbital debris, but there are a number of uncertainties in these models, and additional analysis and data are required to fully characterize the future environment. However, the nature of these uncertainties are such that while the future environment is uncertain, the fact that collisions will control the future environment is less uncertain. The data that already exist is sufficient to show that cascading collisions will control the future debris environment with no, or very minor increases in the current low-earth-orbit population. Two populations control this process: explosion fragments and expended rocket bodies and payloads. Practices are already changing to limit explosions in low earth orbit; it is necessary to begin limiting the number of expended rocket bodies and payloads in orbit.

Computational Aerothermodynamics in Aeroassist Applications

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2001-01-01

Aeroassisted planetary entry uses atmospheric drag to decelerate spacecraft from super-orbital to orbital or suborbital velocities. Numerical simulation of flow fields surrounding these spacecraft during hypersonic atmospheric entry is required to define aerothermal loads. The severe compression in the shock layer in front of the vehicle and subsequent, rapid expansion into the wake are characterized by high temperature, thermo-chemical nonequilibrium processes. Implicit algorithms required for efficient, stable computation of the governing equations involving disparate time scales of convection, diffusion, chemical reactions, and thermal relaxation are discussed. Robust point-implicit strategies are utilized in the initialization phase; less robust but more efficient line-implicit strategies are applied in the endgame. Applications to ballutes (balloon-like decelerators) in the atmospheres of Venus, Mars, Titan, Saturn, and Neptune and a Mars Sample Return Orbiter (MSRO) are featured. Examples are discussed where time-accurate simulation is required to achieve a steady-state solution.

Gate control of quantum dot-based electron spin-orbit qubits

NASA Astrophysics Data System (ADS)

Wu, Shudong; Cheng, Liwen; Yu, Huaguang; Wang, Qiang

2018-07-01

We investigate theoretically the coherent spin dynamics of gate control of quantum dot-based electron spin-orbit qubits subjected to a tilted magnetic field under electric-dipole spin resonance (EDSR). Our results reveal that Rabi oscillation of qubit states can be manipulated electrically based on rapid gate control of SOC strength. The Rabi frequency is strongly dependent on the gate-induced electric field, the strength and orientation of the applied magnetic field. There are two major EDSR mechanisms. One arises from electric field-induced spin-orbit hybridization, and the other arises from magnetic field-induced energy-level crossing. The SOC introduced by the gate-induced electric field allows AC electric fields to drive coherent Rabi oscillations between spin-up and -down states. After the crossing of the energy-levels with the magnetic field, the spin-transfer crossing results in Rabi oscillation irrespective of whether or not the external electric field is present. The spin-orbit qubit is transferred into the orbit qubit. Rabi oscillation is anisotropic and periodic with respect to the tilted and in-plane orientation of the magnetic field originating from the interplay of the SOC, orbital, and Zeeman effects. The strong electrically-controlled SOC strength suggests the possibility for scalable applications of gate-controllable spin-orbit qubits.

Controlling chaos faster.

PubMed

Bick, Christian; Kolodziejski, Christoph; Timme, Marc

2014-09-01

Predictive feedback control is an easy-to-implement method to stabilize unknown unstable periodic orbits in chaotic dynamical systems. Predictive feedback control is severely limited because asymptotic convergence speed decreases with stronger instabilities which in turn are typical for larger target periods, rendering it harder to effectively stabilize periodic orbits of large period. Here, we study stalled chaos control, where the application of control is stalled to make use of the chaotic, uncontrolled dynamics, and introduce an adaptation paradigm to overcome this limitation and speed up convergence. This modified control scheme is not only capable of stabilizing more periodic orbits than the original predictive feedback control but also speeds up convergence for typical chaotic maps, as illustrated in both theory and application. The proposed adaptation scheme provides a way to tune parameters online, yielding a broadly applicable, fast chaos control that converges reliably, even for periodic orbits of large period.

CSM docked DAP/orbital assembly bending interaction-axial case

NASA Technical Reports Server (NTRS)

Turnbull, J. F.; Jones, J. E.

1972-01-01

A digital autopilot which can provide attitude control for the entire Skylab orbital assembly using the service module reaction control jets is described. An important consideration is the potential interaction of the control system with the bending modes of the orbital assembly. Two aspects of this potential interaction were considered. The first was the possibility that bending induced rotations feeding back through the attitude sensor into the control system could produce an instability or self-sustained oscillation. The second was whether the jet activity commanded by the control system could produce excessive loads at any of the critical load points of the orbital assembly. Both aspects were studied by using analytic techniques and by running simulations on the all-digital simulator.

A methodology for obtaining on-orbit SI-traceable spectral radiance measurements in the thermal infrared

NASA Astrophysics Data System (ADS)

Dykema, John A.; Anderson, James G.

2006-06-01

A methodology to achieve spectral thermal radiance measurements from space with demonstrable on-orbit traceability to the International System of Units (SI) is described. This technique results in measurements of infrared spectral radiance R(\\tilde {\\upsilon }) , with spectral index \\tilde {\\upsilon } in cm-1, with a relative combined uncertainty u_c[R(\\tilde {\\upsilon })] of 0.0015 (k = 1) for the average mid-infrared radiance emitted by the Earth. This combined uncertainty, expressed in brightness temperature units, is equivalent to ±0.1 K at 250 K at 750 cm-1. This measurement goal is achieved by utilizing a new method for infrared scale realization combined with an instrument design optimized to minimize component uncertainties and admit tests of radiometric performance. The SI traceability of the instrument scale is established by evaluation against source-based and detector-based infrared scales in defined laboratory protocols before launch. A novel strategy is executed to ensure fidelity of on-orbit calibration to the pre-launch scale. This strategy for on-orbit validation relies on the overdetermination of instrument calibration. The pre-launch calibration against scales derived from physically independent paths to the base SI units provides the foundation for a critical analysis of the overdetermined on-orbit calibration to establish an SI-traceable estimate of the combined measurement uncertainty. Redundant calibration sources and built-in diagnostic tests to assess component measurement uncertainties verify the SI traceability of the instrument calibration over the mission lifetime. This measurement strategy can be realized by a practical instrument, a prototype Fourier-transform spectrometer under development for deployment on a small satellite. The measurement record resulting from the methodology described here meets the observational requirements for climate monitoring and climate model testing and improvement.

Orbiter subsystem hardware/software interaction analysis. Volume 8: Forward reaction control system

NASA Technical Reports Server (NTRS)

Becker, D. D.

1980-01-01

The results of the orbiter hardware/software interaction analysis for the AFT reaction control system are presented. The interaction between hardware failure modes and software are examined in order to identify associated issues and risks. All orbiter subsystems and interfacing program elements which interact with the orbiter computer flight software are analyzed. The failure modes identified in the subsystem/element failure mode and effects analysis are discussed.

Copernicus POD Service: Orbit Determination of the Sentinel Satellites

NASA Astrophysics Data System (ADS)

Peter, Heike; Fernández, Jaime; Ayuga, Francisco; Féménias, Pierre

2016-04-01

The Copernicus POD (Precise Orbit Determination) Service is part of the Copernicus Processing Data Ground Segment (PDGS) of the Sentinel-1, -2 and -3 missions. A GMV-led consortium is operating the Copernicus POD Service being in charge of generating precise orbital products and auxiliary data files for their use as part of the processing chains of the respective Sentinel PDGS. Sentinel-1A was launched in April 2014 while Sentinel-2A was on June 2015 and both are routinely operated since then. Sentinel-3A is expected to be launched in February 2016 and Sentinel-1B is planned for spring 2016. Thus the CPOD Service will be operating three to four satellites simultaneously in spring 2016. The satellites of the Sentinel-1, -2, and -3 missions are all equipped with dual frequency high precision GPS receivers delivering the main observables for POD. Sentinel-3 satellites will additionally be equipped with a laser retro reflector for Satellite Laser Ranging and a receiver for DORIS tracking. All three types of observables (GPS, SLR and DORIS) will be used routinely for POD. The POD core of the CPOD Service is NAPEOS (Navigation Package for Earth Orbiting Satellites) the leading ESA/ESOC software for precise orbit determination. The careful selection of models and inputs is important to achieve the different but very demanding requirements in terms of orbital accuracy and timeliness for the Sentinel -1, -2 & -3 missions. The three missions require orbital products with various latencies from 30 minutes up to 20-30 days. The accuracy requirements are also different and partly very challenging, targeting 5 cm in 3D for Sentinel-1 and 2-3 cm in radial direction for Sentinel-3. Although the characteristics and the requirements are different for the three missions the same core POD setup is used to the largest extent possible. This strategy facilitates maintenance of the complex system of the CPOD Service. Updates in the dynamical modelling of the satellite orbits, e.g. improvements of the box-wing models, have been done to deliver best possible orbit solutions for the satellite. Quality control of the CPOD orbits is done by validating them with independent orbit solutions provided by the Copernicus POD Quality Working Group. The cross-comparison of orbit solutions from different institutions is essential to monitor and to improve the orbit accuracy because for Sentinel-1 and -2 this is the only possibility to externally assess the quality of the orbits. Sentinel-3 orbits may additionally be validated by using SLR and DORIS observations. This paper presents the Copernicus POD Service in terms of operations and orbital accuracy achieved by the different orbit products of the different missions. For Sentinel-1 and Sentinel-2, this paper presents the impact of the box-wing models. For Sentinel-3, the orbital accuracy will be assessed using the very first data after launch.

Relative attitude dynamics and control for a satellite inspection mission

NASA Astrophysics Data System (ADS)

Horri, Nadjim M.; Kristiansen, Kristian U.; Palmer, Phil; Roberts, Mark

2012-02-01

The problem of conducting an inspection mission from a chaser satellite orbiting a target spaceraft is considered. It is assumed that both satellites follow nearly circular orbits. The relative orbital motion is described by the Hill-Clohessy-Wiltshire equation. In the case of an elliptic relative orbit, it is shown that an inspection mission is feasible when the chaser is inertially pointing, provided that the camera mounted on the chaser satellite has sufficiently large field of view. The same possibility is shown when the optical axis of the chaser's camera points in, or opposite to, the tangential direction of the local vertical local horizontal frame. For an arbitrary relative orbit and arbitrary initial conditions, the concept of relative Euler angles is defined for this inspection mission. The expression of the desired relative angular velocity vector is derived as a function of Cartesian coordinates of the relative orbit. A quaternion feedback controller is then designed and shown to perform relative attitude control with admissible internal torques. Three different types of relative orbits are considered, namely the elliptic, Pogo and drifting relative orbits. Measurements of the relative orbital motion are assumed to be available from optical navigation.

Controllability of Large SEP for Earth Orbit Raising

NASA Technical Reports Server (NTRS)

Woodcock, Gordon

2004-01-01

A six-degree-of-freedom (6DOF) simulation was constructed and exercised for a large solar electric propulsion (SEP) vehicle operating in low Earth orbit Nominal power was 500 kWe, with the large array sizes implied. Controllability issues, including gravity gradient, roll maneuvering for Sun tracking, and flexible arrays, and flight control methods, were investigated. Initial findings are that a SEP vehicle of this size is controllable and could be used for orbit raising of heavy payloads.

Sustaining high-energy orbits of bi-stable energy harvesters by attractor selection

NASA Astrophysics Data System (ADS)

Udani, Janav P.; Arrieta, Andres F.

2017-11-01

Nonlinear energy harvesters have the potential to efficiently convert energy over a wide frequency range; however, difficulties in attaining and sustaining high-energy oscillations restrict their applicability in practical scenarios. In this letter, we propose an actuation methodology to switch the state of bi-stable harvesters from the low-energy intra-well configuration to the coexisting high-energy inter-well configuration by controlled phase shift perturbations. The strategy is designed to introduce a change in the system state without creating distinct metastable attractors by exploiting the basins of attraction of the coexisting stable attractors. Experimental results indicate that the proposed switching strategy yields a significant improvement in energy transduction capabilities, is highly economical, enabling the rapid recovery of energy spent in the disturbance, and can be practically implemented with widely used low-strain piezoelectric transducers.

Cassini Spacecraft In-Flight Swap to Backup Attitude Control Thrusters

NASA Technical Reports Server (NTRS)

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.

A Dynamical Systems Approach to the Design of the Science Orbit Around Europa

NASA Technical Reports Server (NTRS)

Gomez, Gerard; Lara, Martin; Russell, Ryan P.

2006-01-01

The science orbit for a future mission to Europa requires low eccentricity, low altitude, and high inclination. However, high inclination orbits around planetary satellites are unstable due to third-body perturbations. Without control, the orbiter impacts Europa after few weeks. To minimize control, a tour over the stable-unstable, averaged manifolds of unstable frozen orbits has been suggested. We proceed with the unaveraged equations and study the manifolds of unstable orbits that are periodic in a rotating frame attached to Europa. Massive numerical computation helps in understanding the unstable dynamics close to Europa, and, thus, in selecting long lifetime high inclination orbits. A final test of a selected set of initial conditions on a high fidelity, ephemeris model, validate the results.

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

NASA Astrophysics Data System (ADS)

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.

Comparison of Ultra-Rapid Orbit Prediction Strategies for GPS, GLONASS, Galileo and BeiDou.

PubMed

Geng, Tao; Zhang, Peng; Wang, Wei; Xie, Xin

2018-02-06

Currently, ultra-rapid orbits play an important role in the high-speed development of global navigation satellite system (GNSS) real-time applications. This contribution focuses on the impact of the fitting arc length of observed orbits and solar radiation pressure (SRP) on the orbit prediction performance for GPS, GLONASS, Galileo and BeiDou. One full year's precise ephemerides during 2015 were used as fitted observed orbits and then as references to be compared with predicted orbits, together with known earth rotation parameters. The full nine-parameter Empirical Center for Orbit Determination in Europe (CODE) Orbit Model (ECOM) and its reduced version were chosen in our study. The arc lengths of observed fitted orbits that showed the smallest weighted root mean squares (WRMSs) and medians of the orbit differences after a Helmert transformation fell between 40 and 45 h for GPS and GLONASS and between 42 and 48 h for Galileo, while the WRMS values and medians become flat after a 42 h arc length for BeiDou. The stability of the Helmert transformation and SRP parameters also confirmed the similar optimal arc lengths. The range around 42-45 h is suggested to be the optimal arc length interval of the fitted observed orbits for the multi-GNSS joint solution of ultra-rapid orbits.

Comparison of Ultra-Rapid Orbit Prediction Strategies for GPS, GLONASS, Galileo and BeiDou

PubMed Central

Zhang, Peng; Wang, Wei; Xie, Xin

2018-01-01

Currently, ultra-rapid orbits play an important role in the high-speed development of global navigation satellite system (GNSS) real-time applications. This contribution focuses on the impact of the fitting arc length of observed orbits and solar radiation pressure (SRP) on the orbit prediction performance for GPS, GLONASS, Galileo and BeiDou. One full year’s precise ephemerides during 2015 were used as fitted observed orbits and then as references to be compared with predicted orbits, together with known earth rotation parameters. The full nine-parameter Empirical Center for Orbit Determination in Europe (CODE) Orbit Model (ECOM) and its reduced version were chosen in our study. The arc lengths of observed fitted orbits that showed the smallest weighted root mean squares (WRMSs) and medians of the orbit differences after a Helmert transformation fell between 40 and 45 h for GPS and GLONASS and between 42 and 48 h for Galileo, while the WRMS values and medians become flat after a 42 h arc length for BeiDou. The stability of the Helmert transformation and SRP parameters also confirmed the similar optimal arc lengths. The range around 42–45 h is suggested to be the optimal arc length interval of the fitted observed orbits for the multi-GNSS joint solution of ultra-rapid orbits. PMID:29415467

Launch Order, Launch Separation, and Loiter in the Constellation 1 1/2-Launch Solution

NASA Technical Reports Server (NTRS)

Stromgren, Chel; Cates, Grant; Cirillo, William

2009-01-01

The NASA Constellation Program (CxP) is developing a two-element Earth-to-Orbit launch system to enable human exploration of the Moon. The first element, Ares I, is a human-rated system that consists of a first stage based on the Space Shuttle Program's solid rocket booster (SRB) and an upper stage that consists of a four-crew Orion capsule, a service module, and a Launch Escape System. The second element, Ares V, is a Saturn V-plus category launch system that consists of the core stage with a cluster of six RS-68B engines and augmented with two 5.5-segment SRBs, a Saturn-derived J-2X engine powering an Earth Departure Stage (EDS), and the lunar-lander vehicle payload, Altair. Initial plans called for the Ares V to be launched first, followed the next day by the Ares I. After the EDS performs the final portion of ascent and subsequent orbit circularization, the Orion spacecraft then performs a rendezvous and docks with the EDS and its Altair payload. Following checkout, the integrated stack loiters in low Earth orbit (LEO) until the appropriate Trans-Lunar Injection (TLI) window opportunity opens, at which time the EDS propels the integrated Orion Altair to the Moon. Successful completion of this 1 1/2-launch solution carries risks related to both the orbital lifetime of the assets and the probability of achieving the launch of the second vehicle within the orbital lifetime of the first. These risks, which are significant in terms of overall system design choices and probability of mission success, dictated a thorough reevaluation of the launch strategy, including the order of vehicle launch and the planned time period between launches. The goal of the effort described in this paper was to select a launch strategy that would result in the greatest possible expected system performance, while accounting for launch risks and the cost of increased orbital lifetime. Discrete Event Simulation (DES) model of the launch strategies was created to determine the probability of a second launch not occurring in a timely fashion (i.e., before the assets waiting in LEO expire). An overview of the launch strategy evaluation process is presented, along with results of specific cases that were analyzed. A high-level comparison of options is then presented, along with the conclusion derived from the analysis.

SPS attitude control and stationkeeping: Requirements and tradeoffs

NASA Technical Reports Server (NTRS)

Oglevie, R. E.

1980-01-01

The dominant control requirements of solar power satellites change appreciably relative to small contemporary spacecraft. Trade studies and analyses illustrated preferred control approaches. It was found that the geosynchronous equatorial orbit is preferred over the alternative orbits considered, that the solar pressure orbit perturbation dominates stationkeeping propulsion requirements and that a combined AC and SK system using ion electric propulsion can satisfy the attitude control requirements. It was also found that control system/structural dynamic interaction stability can be obtained through frequency separation with reasonable structural dynamic requirements and simplify spacecraft design.

Optimal Lorentz-augmented spacecraft formation flying in elliptic orbits

NASA Astrophysics Data System (ADS)

Huang, Xu; Yan, Ye; Zhou, Yang

2015-06-01

An electrostatically charged spacecraft accelerates as it moves through the Earth's magnetic field due to the induced Lorentz force, providing a new means of propellantless electromagnetic propulsion for orbital maneuvers. The feasibility of Lorentz-augmented spacecraft formation flying in elliptic orbits is investigated in this paper. Assuming the Earth's magnetic field as a tilted dipole corotating with Earth, a nonlinear dynamical model that characterizes the orbital motion of Lorentz spacecraft in the vicinity of arbitrary elliptic orbits is developed. To establish a predetermined formation configuration at given terminal time, pseudospectral method is used to solve the optimal open-loop trajectories of hybrid control inputs consisted of Lorentz acceleration and thruster-generated control acceleration. A nontilted dipole model is also introduced to analyze the effect of dipole tilt angle via comparisons with the tilted one. Meanwhile, to guarantee finite-time convergence and system robustness against external perturbations, a continuous fast nonsingular terminal sliding mode controller is designed and the closed-loop system stability is proved by Lyapunov theory. Numerical simulations substantiate the validity of proposed open-loop and closed-loop control schemes, and the results indicate that an almost propellantless formation establishment can be achieved by choosing appropriate objective function in the pseudospectral method. Furthermore, compared to the nonsingular terminal sliding mode controller, the closed-loop controller presents superior convergence rate with only a bit more control effort. And the proposed controller can be applied in other Lorentz-augmented relative orbital control problems.

Orbiter/Space lab momentum management for POP orientations

NASA Technical Reports Server (NTRS)

Cox, J. W.

1974-01-01

An angular momentum management scheme applicable to the orbiter/spacelab is described. The basis of the scheme is to periodically maneuver the vehicle through a small angle thereby using the gravity gradient torque to dump momentum from the control moment gyro (CMG) control system. The orbiter is operated with its principal vehicle axis perpendicular to the orbital plane. Numerous case runs were conducted on the hybrid simulation and representative cases are included.

Open-loop control of quasiperiodic thermoacoustic oscillations

NASA Astrophysics Data System (ADS)

Guan, Yu; Gupta, Vikrant; Kashinath, Karthik; Li, Larry K. B.

2017-11-01

The open-loop application of periodic acoustic forcing has been shown to be a potentially effective strategy for controlling periodic thermoacoustic oscillations, but its effectiveness on aperiodic thermoacoustic oscillations is less clear. In this experimental study, we apply periodic acoustic forcing to a ducted premixed flame oscillating quasiperiodically at two incommensurate natural frequencies, f1 and f2. We find that (i) above a critical forcing amplitude, the system locks into the forcing by oscillating only at the forcing frequency ff, producing a closed periodic orbit in phase space with no evidence of the original T2 torus attractor; (ii) the critical forcing amplitude required for lock-in decreases as ff approaches either f1 or f2, resulting in characteristic ∨-shaped lock-in boundaries around the two natural modes; and (iii) for a wide range of forcing frequencies, the system's oscillation amplitude can be reduced to less than 20% of that of the unforced system. These findings show that the open-loop application of periodic acoustic forcing can be an effective strategy for controlling aperiodic thermoacoustic oscillations. This work was supported by the Research Grants Council of Hong Kong (Project No. 16235716 and 26202815).

Spin-Orbit Coupling Controlled J = 3 / 2 Electronic Ground State in 5 d 3 Oxides

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

Taylor, A. E.; Calder, S.; Morrow, R.

Entanglement of spin and orbital degrees of freedom drives the formation of novel quantum and topological physical states. Here we report resonant inelastic x-ray scattering measurements of the transition metal oxides Ca3LiOsO6 and Ba2YOsO6, which reveals a dramatic spitting of the t2g manifold. We invoke an intermediate coupling approach that incorporates both spin-orbit coupling and electron-electron interactions on an even footing and reveal that the ground state of 5d3-based compounds, which has remained elusive in previously applied models, is a novel spin-orbit entangled J=3/2 electronic ground state. This work reveals the hidden diversity of spin-orbit controlled ground states in 5dmore » systems and introduces a new arena in the search for spin-orbit controlled phases of matter.« less

Optical aperture synthesis: limitations and interest for the earth observation

NASA Astrophysics Data System (ADS)

Brouard, Laurent; Safa, Frederic; Crombez, Vincent; Laubier, David

2017-11-01

For very large telescope diameters, typically above 4 meters, monolithic telescopes can hardly be envisaged for space applications. Optical aperture synthesis can be envisaged in the future for improving the image resolution from high altitude orbits by co-phasing several individual telescopes of smaller size and reconstituting an aperture of large surface. The telescopes can be deployed on a single spacecraft or distributed on several spacecrafts in free flying formation. Several future projects are based on optical aperture synthesis for science or earth observation. This paper specifically discusses the limitations and interest of aperture synthesis technique for Earth observation from high altitude orbits, in particular geostationary orbit. Classical Fizeau and Michelson configurations are recalled, and system design aspects are investigated: synthesis of the Modulation Transfer Function (MTF), integration time and imaging procedure are first discussed then co-phasing strategies and instrument metrology are developed. The discussion is supported by specific designs made at EADS Astrium. As example, a telescope design is presented with a surface of only 6.6 m2 for the primary mirror for an external diameter of 10.6 m allowing a theoretical resolution of 1.2 m from geostationary orbit with a surface lower than 10% of the overall surface. The impact is that the integration time is increasing leading to stringent satellite attitude requirements. Image simulation results are presented. The practical implementation of the concept is evaluated in terms of system impacts in particular spacecraft attitude control, spacecraft operations and imaging capability limitations.

The Earth System Science Pathfinder Orbiting Carbon Observatory (OCO) Mission

NASA Technical Reports Server (NTRS)

Crisp, David

2003-01-01

A viewgraph presentation describing the Earth System Science Pathfinder Orbiting Carbon Observatory (OCO) Mission is shown. The contents include: 1) Why CO2?; 2) What Processes Control CO2 Sinks?; 3) OCO Science Team; 4) Space-Based Measurements of CO2; 5) Driving Requirement: Precise, Bias-Free Global Measurements; 6) Making Precise CO2 Measurements from Space; 7) OCO Spatial Sampling Strategy; 8) OCO Observing Modes; 9) Implementation Approach; 10) The OCO Instrument; 11) The OCO Spacecraft; 12) OCO Will Fly in the A-Train; 13) Validation Program Ensures Accuracy and Minimizes Spatially Coherent Biases; 14) Can OCO Provide the Required Precision?; 15) O2 Column Retrievals with Ground-based FTS; 16) X(sub CO2) Retrieval Simulations; 17) Impact of Albedo and Aerosol Uncertainty on X(sub CO2) Retrievals; 18) Carbon Cycle Modeling Studies: Seasonal Cycle; 19) Carbon Cycle Modeling Studies: The North-South Gradient in CO2; 20) Carbon Cycle Modeling Studies: Effect of Diurnal Biases; 21) Project Status and Schedule; and 22) Summary.

Orbital controlled band gap engineering of tetragonal BiFeO 3 for optoelectronic applications

DOE PAGES

Qiao, L.; Zhang, S.; Xiao, H. Y.; ...

2018-01-01

Bismuth ferrite BiFeO 3 (BFO) is an important ferroelectric material for thin-film optoelectronic sensing and potential photovoltaic applications. Its relatively large band gap, however, limits the conversion efficiency of BFO absorber-based PV devices. In this study, based on density functional theory calculations we demonstrate that with well-designed Fe-site elemental substitution, tetragonal BFO can exhibit a much lower fundamental band gap than conventional rhombohedral BFO without forming in-gap electronic states and unravel the underlying mechanisms. Cation atomic size, electronegativity, and crystallographic symmetry are evidenced as critical parameters to tailor the metal 3d – oxygen 2p orbital interactions and thus intrinsically modifymore » electronic structure, particularly, the shape and character of the valence and conduction band edges. With reduced band gap, improved mobility, and uncompromised ferroelectric and magnetic ground states, the present results provide a new strategy of designing high symmetry BFO for efficient optoelectronic applications.« less

Orbital controlled band gap engineering of tetragonal BiFeO 3 for optoelectronic applications

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

Qiao, L.; Zhang, S.; Xiao, H. Y.

Bismuth ferrite BiFeO 3 (BFO) is an important ferroelectric material for thin-film optoelectronic sensing and potential photovoltaic applications. Its relatively large band gap, however, limits the conversion efficiency of BFO absorber-based PV devices. In this study, based on density functional theory calculations we demonstrate that with well-designed Fe-site elemental substitution, tetragonal BFO can exhibit a much lower fundamental band gap than conventional rhombohedral BFO without forming in-gap electronic states and unravel the underlying mechanisms. Cation atomic size, electronegativity, and crystallographic symmetry are evidenced as critical parameters to tailor the metal 3d – oxygen 2p orbital interactions and thus intrinsically modifymore » electronic structure, particularly, the shape and character of the valence and conduction band edges. With reduced band gap, improved mobility, and uncompromised ferroelectric and magnetic ground states, the present results provide a new strategy of designing high symmetry BFO for efficient optoelectronic applications.« less

Short-Arc Orbit Determination Results and Space Debris Test Observation of the OWL-Net

NASA Astrophysics Data System (ADS)

Choi, J.; Jo, J.; Yim, H.

Korea Astronomy and Space Science Institute had developed the Optical Wide-field patroL-Network (OWL-Net) for maintaining the domestic Low Earth Orbit satellites’ ephemeris and monitoring Geostationary Earth Orbit region. It also can be used to observe space debris. The orbit determination process was planned with batch least square orbit estimator for every week. The optical tracking window is very narrow, several times per week. Sequentialbatch type estimation strategy was attempted for more reliable orbit prediction. We compared the test operation results with Two Line Elements and CPF files to validate the system. This results can be used to estimate the performance of the OWL-Net operations. And also we had observation of the Astro-H debris. We got the dozens of photometric data of the Astro-H debris main part for a few seconds with the chopper system.

Compensator-based 6-DOF control for probe asteroid-orbital-frame hovering with actuator limitations

NASA Astrophysics Data System (ADS)

Liu, Xiaosong; Zhang, Peng; Liu, Keping; Li, Yuanchun

2016-05-01

This paper is concerned with 6-DOF control of a probe hovering in the orbital frame of an asteroid. Considering the requirements of the scientific instruments pointing direction and orbital position in practical missions, the coordinate control of relative attitude and orbit between the probe and target asteroid is imperative. A 6-DOF dynamic equation describing the relative translational and rotational motion of a probe in the asteroid's orbital frame is derived, taking the irregular gravitation, model and parameter uncertainties and external disturbances into account. An adaptive sliding mode controller is employed to guarantee the convergence of the state error, where the adaptation law is used to estimate the unknown upper bound of system uncertainty. Then the controller is improved to deal with the practical problem of actuator limitations by introducing a RBF neural network compensator, which is used to approximate the difference between the actual control with magnitude constraint and the designed nominal control law. The closed-loop system is proved to be asymptotically stable through the Lyapunov stability analysis. Numerical simulations are performed to compare the performances of the preceding designed control laws. Simulation results demonstrate the validity of the control scheme using the compensator-based adaptive sliding mode control law in the presence of actuator limitations, system uncertainty and external disturbance.

Astrodynamics 1991; Proceedings of the AAS/AIAA Astrodynamics Conference, Durango, CO, Aug. 19-22, 1991. Pts. 1, 2, and 3

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

Kaufman, B.; Alfriend, K.T.; Roehrich, R.L.

1992-01-01

The present conference on astrodynamics and advances in the astronautical sciences encompasses orbit determination, orbital debris, flexible-body dynamics and control, attitude dynamics and control, and topics related to the projects of the European space program. Specific issues addressed include a numerical approach to the angles-only initial orbit determination problem, precise orbit determination of the SPOT platform with DORIS, space-debris measurement and modeling, H(infinity)-optimized broadband compensator for wave-absorbing control, and the application of linear actuators for for telescope pointing control. Also addressed are attitude determination and dynamical performance in free drift for the Space Station Freedom, a Kalman filter for amore » gravity-gradient satellite, the positioning of the Eutelsat II satellite from supersynchronous transfer orbit to reduce satellite velocity-correction requirements, and trajectory analysis and issues.« less

Development of preliminary design concept for a multifunction display and control system for the Orbiter crew station. Task 4: Design concept recommendation

NASA Technical Reports Server (NTRS)

Spiger, R. J.; Farrell, R. J.; Holcomb, G. A.

1982-01-01

Application of multifunction display and control systems to the NASA Orbiter spacecraft offers the potential for reducing crew workload and improving the presentation of system status and operational data to the crew. A design concept is presented for the application of a multifunction display and control system (MFDCS) to the Orbital Maneuvering System and Electrical Power Distribution and Control System on the Orbiter spacecraft. The MFDCS would provide the capability for automation of procedures, fault prioritization and software reconfiguration of the MFDCS data base. The MFDCS would operate as a stand-alone processor to minimize the impact on the current Orbiter software. Supervisory crew command of all current functions would be retained through the use of several operating modes in the system. Both the design concept and the processes followed in defining the concept are described.

Electric control of superconducting transition through a spin-orbit coupled interface

PubMed Central

Ouassou, Jabir Ali; Di Bernardo, Angelo; Robinson, Jason W. A.; Linder, Jacob

2016-01-01

We demonstrate theoretically all-electric control of the superconducting transition temperature using a device comprised of a conventional superconductor, a ferromagnetic insulator, and semiconducting layers with intrinsic spin-orbit coupling. By using analytical calculations and numerical simulations, we show that the transition temperature of such a device can be controlled by electric gating which alters the ratio of Rashba to Dresselhaus spin-orbit coupling. The results offer a new pathway to control superconductivity in spintronic devices. PMID:27426887

ATS-6 engineering performance report. Volume 2: Orbit and attitude controls

NASA Technical Reports Server (NTRS)

Wales, R. O. (Editor)

1981-01-01

Attitude control is reviewed, encompassing the attitude control subsystem, spacecraft attitude precision pointing and slewing adaptive control experiment, and RF interferometer experiment. The spacecraft propulsion system (SPS) is discussed, including subsystem, SPS design description and validation, orbital operations and performance, in-orbit anomalies and contingency operations, and the cesium bombardment ion engine experiment. Thruster failure due to plugging of the propellant feed passages, a major cause for mission termination, are considered among the critical generic failures on the satellite.

Design and optimal control of on-orbit servicing trajectory for target vehicle in non-coplanar elliptical orbit

NASA Astrophysics Data System (ADS)

Zhou, Wenyong; Yuan, Jianping; Luo, Jianjun

2005-11-01

Autonomous on-orbit servicing provides flexibility to space systems and has great value both in civil and in military. When a satellite performs on-orbit servicing tasks, flying around is the basic type of motion. This paper is concerned with the design and control problems of a chaser satellite flying around a target spacecraft in non-coplanar elliptical orbit for a long time. At first, a mathematical model used to design a long-term flying around trajectory is presented, which is applicable to the situation that the target spacecraft flies in an elliptical orbit. The conditions of the target at the centre of the flying around path are deduced. Considering the safety and task requirements, a long-term flying around trajectory is designed. Taking into account perturbations and navigation errors which can cause the trajectory unstable and mission impossible, a two-impulse control method is put forward. Genetic algorithm is used to minimize the cost function which considers fuel consumption and bias simultaneously. Some simulation works are carried out and the results indicate the flying around mathematical model and the trajectory control method can be used in the design and control of a long-term flying around trajectory.

Mission Control Center (MCC) System Specification for the Shuttle Orbital Flight Test (OFT) Timeframe

NASA Technical Reports Server (NTRS)

1976-01-01

System specifications to be used by the mission control center (MCC) for the shuttle orbital flight test (OFT) time frame were described. The three support systems discussed are the communication interface system (CIS), the data computation complex (DCC), and the display and control system (DCS), all of which may interfere with, and share processing facilities with other applications processing supporting current MCC programs. The MCC shall provide centralized control of the space shuttle OFT from launch through orbital flight, entry, and landing until the Orbiter comes to a stop on the runway. This control shall include the functions of vehicle management in the area of hardware configuration (verification), flight planning, communication and instrumentation configuration management, trajectory, software and consumables, payloads management, flight safety, and verification of test conditions/environment.

Orbital design strategy for domestic communication satellite systems.

NASA Technical Reports Server (NTRS)

Ramji, S.; Sawitz, P.

1973-01-01

Review of some of the considerations pertinent to efficient orbit utilization in the design of domestic communications satellite systems. A strategy is developed to efficiently locate a heterogeneous system of satellites within the available arc and provide room for future growth. A practical design is illustrated, using a computer simulation model, for the placement of 25 satellites within 73% of the available arc employing frequency and polarization coordination techniques. A number of widely variable factors that influence satellite spacing are examined. These factors include such critical system elements as telephony and television interference noise limits, frequency plan coordination, polarization plan coordination, ground antenna diameter, signal protection ratio, and satellite station keeping.

Mars Reconnaissance Orbiter Navigation Strategy for the Comet Siding Spring Encounter

NASA Technical Reports Server (NTRS)

Menon, Premkumar R.; Wagner, Sean V.; Martin-Mur, Tomas J.; Jefferson, David C.; Ardalan, Shadan M.; Chung, Min-Kun J.; Lee, Kyong J.; Schulze, William B.

2015-01-01

Comet Siding Spring encountered Mars on October 19, 2014 at a distance of about 140,500 km - the nearest comet flyby of a planet in recorded history. Mars Reconnaissance Orbiter (MRO) was able to detect the comet, gather science data, and capture images of the comet as it approached Mars. To help protect MRO from the incoming comet particles, two propulsive maneuvers were performed to position the spacecraft behind Mars at the arrival time of the expected peak particle fluency. This paper documents the strategy that the MRO Navigation Team executed to mitigate risk from the comet particles while allowing scientific observations of the comet flyby.

Earth Science System of the Future: Observing, Processing, and Delivering Data Products Directly to Users

NASA Technical Reports Server (NTRS)

Crisp, David; Komar, George (Technical Monitor)

2001-01-01

Advancement of our predictive capabilities will require new scientific knowledge, improvement of our modeling capabilities, and new observation strategies to generate the complex data sets needed by coupled modeling networks. New observation strategies must support remote sensing from a variety of vantage points and will include "sensorwebs" of small satellites in low Earth orbit, large aperture sensors in Geostationary orbits, and sentinel satellites at L1 and L2 to provide day/night views of the entire globe. Onboard data processing and high speed computing and communications will enable near real-time tailoring and delivery of information products (i.e., predictions) directly to users.

Exoplanets and SETI

NASA Astrophysics Data System (ADS)

Wright, Jason T.

The discovery of exoplanets has both focused and expanded the search for extraterrestrial intelligence. The consideration of Earth as an exoplanet, the knowledge of the orbital parameters of individual exoplanets, and our new understanding of the prevalence of exoplanets throughout the galaxy have all altered the search strategies of communication SETI efforts, by inspiring new "Schelling points" (i.e. optimal search strategies for beacons). Future efforts to characterize individual planets photometrically and spectroscopically, with imaging and via transit, will also allow for searches for a variety of technosignatures on their surfaces, in their atmospheres, and in orbit around them. In the near-term, searches for new planetary systems might even turn up free-floating megastructures.

NASA's Orbital Space Plane Risk Reduction Strategy

NASA Technical Reports Server (NTRS)

Dumbacher, Dan

2003-01-01

This paper documents the transformation of NASA s Space Launch Initiative (SLI) Second Generation Reusable Launch Vehicle Program under the revised Integrated Space Transportation Plan, announced November 2002. Outlining the technology development approach followed by the original SLI, this paper gives insight into the current risk-reduction strategy that will enable confident development of the Nation s first orbital space plane (OSP). The OSP will perform an astronaut and contingency cargo transportation function, with an early crew rescue capability, thus enabling increased crew size and enhanced science operations aboard the International Space Station. The OSP design chosen for full-scale development will take advantage of the latest innovations American industry has to offer. The OSP Program identifies critical technologies that must be advanced to field a safe, reliable, affordable space transportation system for U.S. access to the Station and low-Earth orbit. OSP flight demonstrators will test crew safety features, validate autonomous operations, and mature thermal protection systems. Additional enabling technologies may be identified during the OSP design process as part of an overall risk-management strategy. The OSP Program uses a comprehensive and evolutionary systems acquisition approach, while applying appropriate lessons learned.

Design of relative trajectories for in orbit proximity operations

NASA Astrophysics Data System (ADS)

Opromolla, Roberto; Fasano, Giancarmine; Rufino, Giancarlo; Grassi, Michele

2018-04-01

This paper presents an innovative approach to design relative trajectories suitable for close-proximity operations in orbit, by assigning high-level constraints regarding their stability, shape and orientation. Specifically, this work is relevant to space mission scenarios, e.g. formation flying, on-orbit servicing, and active debris removal, which involve either the presence of two spacecraft carrying out coordinated maneuvers, or a servicing/recovery spacecraft (chaser) performing monitoring, rendezvous and docking with respect to another space object (target). In the above-mentioned scenarios, an important aspect is the capability of reducing collision risks and of providing robust and accurate relative navigation solutions. To this aim, the proposed approach exploits a relative motion model relevant to two-satellite formations, and developed in mean orbit parameters, which takes the perturbation effect due to secular Earth oblateness, as well as the motion of the target along a small-eccentricity orbit, into account. This model is used to design trajectories which ensure safe relative motion, to minimize collision risks and relax control requirements, providing at the same time favorable conditions, in terms of target-chaser relative observation geometry for pose determination and relative navigation with passive or active electro-optical sensors on board the chaser. Specifically, three design strategies are proposed in the context of a space target monitoring scenario, considering as design cases both operational spacecraft and debris, characterized by highly variable shape, size and absolute rotational dynamics. The effectiveness of the proposed design approach in providing favorable observation conditions for target-chaser relative pose estimation is demonstrated within a simulation environment which reproduces the designed target-chaser relative trajectory, the operation of an active LIDAR installed on board the chaser, and pose estimation algorithms.

Precision GPS orbit determination strategies for an earth orbiter and geodetic tracking system

NASA Technical Reports Server (NTRS)

Lichten, Stephen M.; Bertiger, Willy I.; Border, James S.

1988-01-01

Data from two 1985 GPS field tests were processed and precise GPS orbits were determined. With a combined carrier phase and pseudorange, the 1314-km repeatability improves substantially to 5 parts in 10 to the 9th (0.6 cm) in the north and 2 parts in 10 to the 8th (2-3 cm) in the other components. To achieve these levels of repeatability and accuracy, it is necessary to fine-tune the GPS solar radiation coefficients and ground station zenith tropospheric delays.

On-orbit evaluation of the control system/structural mode interactions on OSO-8

NASA Technical Reports Server (NTRS)

Slafer, L. I.

1980-01-01

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. This 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. The test results have led to a validation of both the analytical modeling and servo design techniques used during the development of the control system, and also verified the approach taken to vehicle and servo ground testing.

Optimization of Planet Finder Observing Strategy

NASA Astrophysics Data System (ADS)

Sinukoff, E.

2014-03-01

We evaluate radial velocity observing strategies to be considered for future planethunting surveys with the Automated Planet Finder, a new 2.4-m telescope at Lick Observatory. Observing strategies can be optimized to mitigate stellar noise, which can mask and imitate the weak Doppler signals of low-mass planets. We estimate and compare sensitivities of 5 different observing strategies to planets around G2-M2 dwarfs, constructing RV noise models for each stellar spectral type, accounting for acoustic, granulation, and magnetic activity modes. The strategies differ in exposure time, nightly and monthly cadence, and number of years. Synthetic RV time-series are produced by injecting a planet signal onto the stellar noise, sampled according to each observing strategy. For each star and each observing strategy, thousands of planet injection recovery trials are conducted to determine the detection efficiency as a function of orbital period, minimum mass, and eccentricity. We find that 4-year observing strategies of 10 nights per month are sensitive to planets ~25-40% lower in mass than the corresponding 1 year strategies of 30 nights per month. Three 5-minute exposures spaced evenly throughout each night provide a 10% gain in sensitivity over the corresponding single 15-minute exposure strategies. All strategies are sensitive to planets of lowest mass around the modeled K7 dwarf. This study indicates that APF surveys adopting the 4-year strategies should detect Earth-mass planets on < 10-day orbits around quiet late-K dwarfs as well as > 1.6 Earth-mass planets in their habitable zones.

STS-49 Endeavour, Orbiter Vehicle (OV) 105, Orbit Team O1 in MCC Bldg 30 FCR

NASA Technical Reports Server (NTRS)

1992-01-01

STS-49 Endeavour, Orbiter Vehicle (OV) 105, Orbit Team 1 (O1) poses in front of large display screens in JSC's Mission Control Center (MCC) Bldg 30 Flight Control Room (FCR) for group portrait. Lead Flight Director (FD) Granvil A. Pennington stands next to a model of the James Cook's ship, the Endeavour (left). Astronaut and Spacecraft Communicator (CAPCOM) John H. Casper stands at the right of the model.

Surface hopping trajectory simulations with spin-orbit and dynamical couplings

NASA Astrophysics Data System (ADS)

Granucci, Giovanni; Persico, Maurizio; Spighi, Gloria

2012-12-01

In this paper we consider the inclusion of the spin-orbit interaction in surface hopping molecular dynamics simulations to take into account spin forbidden transitions. Two alternative approaches are examined. The spin-diabatic one makes use of eigenstates of the spin-free electronic Hamiltonian and of hat{S}^2 and is commonly applied when the spin-orbit coupling is weak. We point out some inconsistencies of this approach, especially important when more than two spin multiplets are coupled. The spin-adiabatic approach is based on the eigenstates of the total electronic Hamiltonian including the spin-orbit coupling. Advantages and drawbacks of both strategies are discussed and illustrated with the help of two model systems.

HERRO Mission to Mars Using Telerobotic Surface Exploration from Orbit

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; Landis, Geoffrey A.; McGuire, Melissa L.; Schmidt, George R.

2013-01-01

This paper presents a concept for a human mission to Mars orbit that features direct robotic exploration of the planet s surface via teleoperation from orbit. This mission is a good example of Human Exploration using Real-time Robotic Operations (HERRO), an exploration strategy that refrains from sending humans to the surfaces of planets with large gravity wells. HERRO avoids the need for complex and expensive man-rated lander/ascent vehicles and surface systems. Additionally, the humans are close enough to the surface to effectively eliminate the two-way communication latency that constrains typical robotic space missions, thus allowing real-time command and control of surface operations and experiments by the crew. Through use of state-of-the-art telecommunications and robotics, HERRO provides the cognitive and decision-making advantages of having humans at the site of study for only a fraction of the cost of conventional human surface missions. It is very similar to how oceanographers and oil companies use telerobotic submersibles to work in inaccessible areas of the ocean, and represents a more expedient, near-term step prior to landing humans on Mars and other large planetary bodies. Results suggest that a single HERRO mission with six crew members could achieve the same exploratory and scientific return as three conventional crewed missions to the Mars surface.

Liquid Acquisition Strategies for Exploration Missions: Current Status 2010

NASA Technical Reports Server (NTRS)

Chato, David J.

2010-01-01

NASA is currently developing the propulsion system concepts for human exploration missions to the lunar surface. The propulsion concepts being investigated are considering the use of cryogenic propellants for the low gravity portion of the mission, that is, the lunar transit, lunar orbit insertion, lunar descent and the rendezvous in lunar orbit with a service module after ascent from the lunar surface. These propulsion concepts will require the vapor free delivery of the cryogenic propellants stored in the propulsion tanks to the exploration vehicles main propulsion system (MPS) engines and reaction control system (RCS) engines. Propellant management devices (PMD s) such as screen channel capillary liquid acquisition devices (LAD s), vanes and sponges currently are used for earth storable propellants in the Space Shuttle Orbiter OMS and RCS applications and spacecraft propulsion applications but only very limited propellant management capability exists for cryogenic propellants. NASA has begun a technology program to develop LAD cryogenic fluid management (CFM) technology through a government in-house ground test program of accurately measuring the bubble point delta-pressure for typical screen samples using LO2, LN2, LH2 and LCH4 as test fluids at various fluid temperatures and pressures. This presentation will document the CFM project s progress to date in concept designs, as well ground testing results.

Dysthyroid Optic Neuropathy.

PubMed

Saeed, Peerooz; Tavakoli Rad, Shahzad; Bisschop, Peter H L T

2018-06-19

Dysthyroid optic neuropathy (DON) is a serious complication of Graves orbitopathy that can result in irreversible and profound visual loss. Controversy exists regarding the pathogenesis and management of the disease. The authors provide an overview of the current understanding of DON and present a therapeutic guideline. A review of the literature. The mechanism of DON appears to be multifactorial: direct compression of the optic nerve by enlarged extraocular muscles, stretching of the optic nerve by proptosis, orbital pressure, vascular insufficiency, and inflammation. Some or all of these factors may be involved in an individual patient. There has only been one controlled trial comparing high-dose intravenous methylprednisolone to bony orbital decompression for DON. Both 2-wall and 3-wall decompression techniques successfully improve visual functions of patients with DON. There are few case reports/case series that suggest biologic agents may improve visual function in DON. DON is a serious complication of Graves orbitopathy, the diagnosis and management of which is complex and requires a multidisciplinary approach. There is little evidence regarding the optimum management strategy. Based on the current literature, the first line of treatment is intravenous methylprednisolone, with the exact timing and indication of bony orbital decompression still to be determined. In addition, there may be a role for the use of biologic agents that will require a systematic program to determine efficacy.

Control of electron spin and orbital resonances in quantum dots through spin-orbit interactions

NASA Astrophysics Data System (ADS)

Stano, Peter; Fabian, Jaroslav

2008-01-01

The influence of a resonant oscillating electromagnetic field on a single electron in coupled lateral quantum dots in the presence of phonon-induced relaxation and decoherence is investigated. Using symmetry arguments, it is shown that the spin and orbital resonances can be efficiently controlled by spin-orbit interactions. The control is possible due to the strong sensitivity of the Rabi frequency to the dot configuration (the orientation of the dot and the applied static magnetic field); the sensitivity is a result of the anisotropy of the spin-orbit interactions. The so-called easy passage configuration is shown to be particularly suitable for a magnetic manipulation of spin qubits, ensuring long spin relaxation times and protecting the spin qubits from electric field disturbances accompanying on-chip manipulations.

Modeling low-thrust transfers between periodic orbits about five libration points: Manifolds and hierarchical design

NASA Astrophysics Data System (ADS)

Zeng, Hao; Zhang, Jingrui

2018-04-01

The low-thrust version of the fuel-optimal transfers between periodic orbits with different energies in the vicinity of five libration points is exploited deeply in the Circular Restricted Three-Body Problem. Indirect optimization technique incorporated with constraint gradients is employed to further improve the computational efficiency and accuracy of the algorithm. The required optimal thrust magnitude and direction can be determined to create the bridging trajectory that connects the invariant manifolds. A hierarchical design strategy dividing the constraint set is proposed to seek the optimal solution when the problem cannot be solved directly. Meanwhile, the solution procedure and the value ranges of used variables are summarized. To highlight the effectivity of the transfer scheme and aim at different types of libration point orbits, transfer trajectories between some sample orbits, including Lyapunov orbits, planar orbits, halo orbits, axial orbits, vertical orbits and butterfly orbits for collinear and triangular libration points, are investigated with various time of flight. Numerical results show that the fuel consumption varies from a few kilograms to tens of kilograms, related to the locations and the types of mission orbits as well as the corresponding invariant manifold structures, and indicates that the low-thrust transfers may be a beneficial option for the extended science missions around different libration points.

A simulation study of emergency lunar escape to orbit using several simplified manual guidance and control techniques

NASA Technical Reports Server (NTRS)

Middleton, D. B.; Hurt, G. J., Jr.

1971-01-01

A fixed-base piloted simulator investigation has been made of the feasibility of using any of several manual guidance and control techniques for emergency lunar escape to orbit with very simplified, lightweight vehicle systems. The escape-to-orbit vehicles accommodate two men, but one man performs all of the guidance and control functions. Three basic attitude-control modes and four manually executed trajectory-guidance schemes were used successfully during approximately 125 simulated flights under a variety of conditions. These conditions included thrust misalinement, uneven propellant drain, and a vehicle moment-of-inertia range of 250 to 12,000 slugs per square foot. Two types of results are presented - orbit characteristics and pilot ratings of vehicle handling qualities.

Free Enterprise: Contributions of the Approach and Landing Test (ALT) Program to the Development of the Space Shuttle Orbiter

NASA Technical Reports Server (NTRS)

Merlin, Peter W.

2006-01-01

The space shuttle orbiter was the first spacecraft designed with the aerodynamic characteristics and in-atmosphere handling qualities of a conventional airplane. In order to evaluate the orbiter's flight control systems and subsonic handling characteristics, a series of flight tests were undertaken at NASA Dryden Flight Research Center in 1977. A modified Boeing 747 Shuttle Carrier Aircraft carried the Enterprise, a prototype orbiter, during eight captive tests to determine how well the two vehicles flew together and to test some of the orbiter s systems. The free-flight phase of the ALT program allowed shuttle pilots to explore the orbiter's low-speed flight and landing characteristics. The Enterprise provided realistic, in-flight simulations of how subsequent space shuttles would be flown at the end of an orbital mission. The fifth free flight, with the Enterprise landing on a concrete runway for the first time, revealed a problem with the space shuttle flight control system that made it susceptible to pilot-induced oscillation, a potentially dangerous control problem. Further research using various aircraft, particularly NASA Dryden's F-8 Digital-Fly-By-Wire testbed, led to correction of the problem before the first Orbital Test Flight.

Close up view of the Commander's Seat on the Flight ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close up view of the Commander's Seat on the Flight Deck of the Orbiter Discovery. Toward the right of the view and in front of te seat is the commander's Rotational Hand Controller. The pilot station has an identical controller. These control the acceleration in the roll pitch and yaw directions via the reaction control system and/or the orbiter maneuvering system while outside of Earth's atmosphere or via the orbiter's aerosurfaces wile in Earth's atmosphere when the atmospheric density permits the surfaces to be effective. There are a number of switches on the controller, most notably a trigger switch which is a push-to-talk switch for voice communication and a large button on top of the controller which is a switch to engage the backup flight system. This view was taken at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

STS-125 Flight Control Team in BFCR - HST Orbit & Planning Teams

NASA Image and Video Library

2009-05-18

JSC2009-E-120479 (18 May 2009) --- Members of the STS-125 Hubble Space Telescope Planning and Orbit flight control team pose for a group portrait in the blue flight control room in the Mission Control Center at NASA's Johnson Space Center.

STS-125 Flight Control Team in BFCR - HST Planning & Orbit Team

NASA Image and Video Library

2009-05-19

JSC2009-E-120701 (19 May 2009) --- Members of the STS-125 Hubble Space Telescope Planning and Orbit flight control team pose for a group portrait in the blue flight control room in the Mission Control Center at NASA's Johnson Space Center.

Three-axis attitude control by two-step rotations using only magnetic torquers in a low Earth orbit near the magnetic equator

NASA Astrophysics Data System (ADS)

Inamori, Takaya; Otsuki, Kensuke; Sugawara, Yoshiki; Saisutjarit, Phongsatorn; Nakasuka, Shinichi

2016-11-01

This study proposes a novel method for three-axis attitude control using only magnetic torquers (MTQs). Previously, MTQs have been utilized for attitude control in many low Earth orbit satellites. Although MTQs are useful for achieving attitude control at low cost and high reliability without the need for propellant, these electromagnetic coils cannot be used to generate an attitude control torque about the geomagnetic field vector. Thus, conventional attitude control methods using MTQs assume the magnetic field changes in an orbital period so that the satellite can generate a required attitude control torque after waiting for a change in the magnetic field direction. However, in a near magnetic equatorial orbit, the magnetic field does not change in an inertial reference frame. Thus, satellites cannot generate a required attitude control torque in a single orbital period with only MTQs. This study proposes a method for achieving a rotation about the geomagnetic field vector by generating a torque that is perpendicular to it. First, this study shows that the three-axis attitude control using only MTQs is feasible with a two-step rotation. Then, the study proposes a method for controlling the attitude with the two-step rotation using a PD controller. Finally, the proposed method is assessed by examining the results of numerical simulations.

Low-cost autonomous orbit control about Mars: Initial simulation results

NASA Astrophysics Data System (ADS)

Dawson, S. D.; Early, L. W.; Potterveld, C. W.; Königsmann, H. J.

1999-11-01

Interest in studying the possibility of extraterrestrial life has led to the re-emergence of the Red Planet as a major target of planetary exploration. Currently proposed missions in the post-2000 period are routinely calling for rendezvous with ascent craft, long-term orbiting of, and sample-return from Mars. Such missions would benefit greatly from autonomous orbit control as a means to reduce operations costs and enable contact with Mars ground stations out of view of the Earth. This paper present results from initial simulations of autonomously controlled orbits around Mars, and points out possible uses of the technology and areas of routine Mars operations where such cost-conscious and robust autonomy could prove most effective. These simulations have validated the approach and control philosophies used in the development of this autonomous orbit controller. Future work will refine the controller, accounting for systematic and random errors in the navigation of the spacecraft from the sensor suite, and will produce prototype flight code for inclusion on future missions. A modified version of Microcosm's commercially available High Precision Orbit Propagator (HPOP) was used in the preparation of these results due to its high accuracy and speed of operation. Control laws were developed to allow an autonomously controlled spacecraft to continuously control to a pre-defined orbit about Mars with near-optimal propellant usage. The control laws were implemented as an adjunct to HPOP. The GSFC-produced 50 × 50 field model of the Martian gravitational potential was used in all simulations. The Martian atmospheric drag was modeled using an exponentially decaying atmosphere based on data from the Mars-GRAM NASA Ames model. It is hoped that the simple atmosphere model that was implemented can be significantly improved in the future so as to approach the fidelity of the Mars-GRAM model in its predictions of atmospheric density at orbital altitudes. Such additional work would take the form of solar flux (F10.7) and diurnal density dependencies. The autonomous controller is a-derivative of the proprietary and patented Microcosm Earth-orbiting control methodology which will be implemented on the upcoming Surrey Satellite Technology (SSTL) UoSAT-12 and the NASA EO-1 spacecraft missions. This work was funded by the NASA Jet Propulsion Laboratory under a Phase I SBIR (96.1 07.02 9444) and by internal Microcosm R&D funds as well as earlier supporting work done under a variety of USAF Research Laboratory-sponsored contracts [1, 2, 4, 12].

[The use of titanium nickelide devices in treating fractures of the zygomatico-orbital complex].

PubMed

Medvedev, Iu A; Sivolapov, K A

1993-01-01

The authors analyze the results of surgical treatment of 78 patients with zygomatico-orbital injuries. Mini-cramps and mini-plates of titanium nickelide with thermomechanical memory were employed for fracture fixation. Surgical strategy based on osteosynthesis with the use of such devices provides a reliable fixation of bone fragments and makes the operation less traumatic.

Countering the Chinese Threat to Low Earth Orbit Satellites: Building a Defensive Space Strategy

DTIC Science & Technology

2009-01-01

into, range of its direct~: . ascent capability. Co-orbital weapons can also be launched fromahostspacecraft,l~etheB~- r on Shenzhou-7. The United States...channel= awst &id=news/awO122 07p2.xml (accessed December 21,2008). DeBlois, Bruce M., ed. Beyond the Paths ofHeaven: The Emergence ofSpace Power Thought

Fuel-optimal, low-thrust transfers between libration point orbits

NASA Astrophysics Data System (ADS)

Stuart, Jeffrey R.

Mission design requires the efficient management of spacecraft fuel to reduce mission cost, increase payload mass, and extend mission life. High efficiency, low-thrust propulsion devices potentially offer significant propellant reductions. Periodic orbits that exist in a multi-body regime and low-thrust transfers between these orbits can be applied in many potential mission scenarios, including scientific observation and communications missions as well as cargo transport. In light of the recent discovery of water ice in lunar craters, libration point orbits that support human missions within the Earth-Moon region are of particular interest. This investigation considers orbit transfer trajectories generated by a variable specific impulse, low-thrust engine with a primer-vector-based, fuel-optimizing transfer strategy. A multiple shooting procedure with analytical gradients yields rapid solutions and serves as the basis for an investigation into the trade space between flight time and consumption of fuel mass. Path and performance constraints can be included at node points along any thrust arc. Integration of invariant manifolds into the design strategy may also yield improved performance and greater fuel savings. The resultant transfers offer insight into the performance of the variable specific impulse engine and suggest novel implementations of conventional impulsive thrusters. Transfers incorporating invariant manifolds demonstrate the fuel savings and expand the mission design capabilities that are gained by exploiting system symmetry. A number of design applications are generated.

State-dependent impulsive models of integrated pest management (IPM) strategies and their dynamic consequences.

PubMed

Tang, Sanyi; Cheke, Robert A

2005-03-01

A state-dependent impulsive model is proposed for integrated pest management (IPM). IPM involves combining biological, mechanical, and chemical tactics to reduce pest numbers to tolerable levels after a pest population has reached its economic threshold (ET). The complete expression of an orbitally asymptotically stable periodic solution to the model with a maximum value no larger than the given ET is presented, the existence of which implies that pests can be controlled at or below their ET levels. We also prove that there is no periodic solution with order larger than or equal to three, except for one special case, by using the properties of the LambertW function and Poincare map. Moreover, we show that the existence of an order two periodic solution implies the existence of an order one periodic solution. Various positive invariant sets and attractors of this impulsive semi-dynamical system are described and discussed. In particular, several horseshoe-like attractors, whose interiors can simultaneously contain stable order 1 periodic solutions and order 2 periodic solutions, are found and the interior structure of the horseshoe-like attractors is discussed. Finally, the largest invariant set and the sufficient conditions which guarantee the global orbital and asymptotic stability of the order 1 periodic solution in the meaningful domain for the system are given using the Lyapunov function. Our results show that, in theory, a pest can be controlled such that its population size is no larger than its ET by applying effects impulsively once, twice, or at most, a finite number of times, or according to a periodic regime. Moreover, our theoretical work suggests how IPM strategies could be used to alter the levels of the ET in the farmers' favour.

Station-Keeping Strategies for Lead-Trail Formation Flying

NASA Astrophysics Data System (ADS)

Martinot, V.; Rozanes, P.

Numerous projects in the Science and Observation domains involve the use of formation flying to ensure the mission performance. The formation flying configurations proposed in some of them are quite complex with several satellites in different planes generating relative differential motions between the satellites like in case of circular projected formation-flying. However, more simple designs consisting of two satellites in a lead-trail formation appear to be sufficient for a wide range of applications (interferometry, geodesy,...). This article concentrates on the station- keeping phase of such formations in Low-Earth Orbits The station-keeping criterion for such formations can be expressed for example in terms of difference in argument of latitude between both satellites and at the altitudes considered, it evolves mainly under the differential effect of the atmospheric drag between the trailing and leading satellites. In the present paper, this differential effect is supposed to originate from the difference in the area-to-mass ratio between the satellites due to their different designs. A preliminary estimation of the navigation performance is first given assuming that on-board GPS receiver are mounted on each satellite of the formation to acquire pseudo-range measurements between the LEO satellites and the MEO GPS constellation. The distance between both satellites of the formation is derived from independent orbit restitution performed for each LEO satellite in a ground master control station processing the GPS measurements. A strategy for controlling the satellite formation disturbed by the differential effect of the drag is then proposed. Simulations are performed to assess the feasibility of the station-keeping with different types of engines. As by-products, the propellant budget and the frequency of the station-keeping manoeuvres are also given. A case study inspired from the ESA project Acechem/Metop is used for the simulations.

Cryogenic thermal system analysis for orbital propellant depot

NASA Astrophysics Data System (ADS)

Chai, Patrick R.; Wilhite, Alan W.

2014-09-01

In any manned mission architecture, upwards of seventy percent of all payload delivered to orbit is propellant, and propellant mass fraction dominates almost all transportation segments of any mission requiring a heavy lift launch system like the Saturn V. To mitigate this, the use of an orbital propellant depot has been extensively studied. In this paper, a thermal model of an orbital propellant depot is used to examine the effects of passive and active thermal management strategies. Results show that an all passive thermal management strategy results in significant boil-off for both hydrogen and oxygen. At current launch vehicle prices, these boil-offs equate to millions of dollars lost per month. Zero boil-off of propellant is achievable with the use of active cryocoolers; however, the cooling power required to produce zero-boil-off is an order of magnitude higher than current state-of-the-art cryocoolers. This study shows a zero-boil-off cryocooler minimum power requirement of 80-100 W at 80 K for liquid oxygen, and 100-120 W at 20 K for liquid hydrogen for a representative Near-Earth Object mission. Research and development effort is required to improve the state-of-the-arts in-space cryogenic thermal management.

Reactive underwater object inspection based on artificial electric sense.

PubMed

Lebastard, Vincent; Boyer, Frédéric; Lanneau, Sylvain

2016-07-26

Weakly electric fish can perform complex cognitive tasks based on extracting information from blurry electric images projected from their immediate environment onto their electro-sensitive skin. In particular they can be trained to recognize the intrinsic properties of objects such as their shape, size and electric nature. They do this by means of novel perceptual strategies that exploit the relations between the physics of a self-generated electric field, their body morphology and the ability to perform specific movement termed probing motor acts (PMAs). In this article we artificially reproduce and combine these PMAs to build an autonomous control strategy that allows an artificial electric sensor to find electrically contrasted objects, and to orbit around them based on a minimum set of measurements and simple reactive feedback control laws of the probe's motion. The approach does not require any simulation models and could be implemented on an autonomous underwater vehicle (AUV) equipped with artificial electric sense. The AUV has only to satisfy certain simple geometric properties, such as bi-laterally (left/right) symmetrical electrodes and possess a reasonably high aspect (length/width) ratio.

ESA technology flies on Italian mini-satellite launched from Russia

NASA Astrophysics Data System (ADS)

2000-07-01

Owned by the Italian space agency (ASI) and developed by Carlo Gavazzi with contributions from many other Italian companies, MITA has two tasks to perform: in a circular orbit at 450 km altitude, the mini satellite will carry a cosmic particle detector, while its platform will be tested for the first time as a vehicle for future scientific missions. MITA also carries the MTS-AOMS payload (MicroTechSensor for Attitude and Orbit Measurement System), developed by Astrium in the framework of ESA's Technology Flight Opportunity trial programme. With the Technology Flight Opportunity scheme, funded by its General Studies Programme, ESA intends to provide access to space for European industry's technology products needing in-orbit demonstration to enhance their competitiveness on the space market. This new form of support to the European space industry ties in with ESA's strategy for fostering the competitiveness of European-made technology for eventual commercialisation. In-orbit demonstration is essential if new technologies are to compete on level terms on non-European markets. It thus consolidates strategic investments made by the space industry. The MTS-AOMS is a highly integrated sensor for autonomous attitude and orbit control systems. It combines three functions in one unit: Earth sensing, star sensing and magnetic field sensing. The equipment incorporates an active pixel array sensor and a 2-D fluxgate magnetometer. The aims of the flight are to verify in situ the payload's inherent functions and performance, which cannot be done on the ground, and to assess the behaviour of this type of technology when exposed to the space environment. The Technology Flight Opportunity rule is that ESA funds the launch and integration costs, industry the development and operating costs. According to present planning, two further in-orbit demonstrations funded by this scheme will be carried out between now and January 2001.

Spacecraft formation control using analytical finite-duration approaches

NASA Astrophysics Data System (ADS)

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.

Operational factors affecting microgravity levels in orbit

NASA Technical Reports Server (NTRS)

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.

Otolith and Vertical Canal Contributions to Dynamic Postural Control

NASA Technical Reports Server (NTRS)

Black, F. Owen

1999-01-01

The objective of this project is to determine: 1) how do normal subjects adjust postural movements in response to changing or altered otolith input, for example, due to aging? and 2) how do patients adapt postural control after altered unilateral or bilateral vestibular sensory inputs such as ablative inner ear surgery or ototoxicity, respectively? The following hypotheses are under investigation: 1) selective alteration of otolith input or abnormalities of otolith receptor function will result in distinctive spatial, frequency, and temporal patterns of head movements and body postural sway dynamics. 2) subjects with reduced, altered, or absent vertical semicircular canal receptor sensitivity but normal otolith receptor function or vice versa, should show predictable alterations of body and head movement strategies essential for the control of postural sway and movement. The effect of altered postural movement control upon compensation and/or adaptation will be determined. These experiments provide data for the development of computational models of postural control in normals, vestibular deficient subjects and normal humans exposed to unusual force environments, including orbital space flight.

STS-106 Orbit 2 Flight Team

NASA Image and Video Library

2000-09-14

JSC2000-06244 (September 2000)--- Flight director Jeff Hanley, front center, and the fifty-odd flight controllers making up the ISS Orbit 2 Team pose for their group portrait in the ISS Flight Control Room of Houston's Mission Control Center.

Orbiter active thermal control system description

NASA Technical Reports Server (NTRS)

Laubach, G. E.

1975-01-01

A brief description of the Orbiter Active Thermal Control System (ATCS) including (1) major functional requirements of heat load, temperature control and heat sink utilization, (2) the overall system arrangement, and (3) detailed description of the elements of the ATCS.

STS-120 Orbit 2 Flight Control Team Photo

NASA Image and Video Library

2007-10-31

JSC2007-E-095908 (31 Oct. 2007) --- The members of the STS-120 Orbit 2 flight control team pose for a group portrait in the space shuttle flight control room of Houston's Mission Control Center (MCC). Flight director Mike Moses holds the STS-120 mission logo.

WE-EF-207-01: FEATURED PRESENTATION and BEST IN PHYSICS (IMAGING): Task-Driven Imaging for Cone-Beam CT in Interventional Guidance

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

Gang, G; Stayman, J; Ouadah, S

2015-06-15

Purpose: This work introduces a task-driven imaging framework that utilizes a patient-specific anatomical model, mathematical definition of the imaging task, and a model of the imaging system to prospectively design acquisition and reconstruction techniques that maximize task-based imaging performance. Utility of the framework is demonstrated in the joint optimization of tube current modulation and view-dependent reconstruction kernel in filtered-backprojection reconstruction and non-circular orbit design in model-based reconstruction. Methods: The system model is based on a cascaded systems analysis of cone-beam CT capable of predicting the spatially varying noise and resolution characteristics as a function of the anatomical model and amore » wide range of imaging parameters. Detectability index for a non-prewhitening observer model is used as the objective function in a task-driven optimization. The combination of tube current and reconstruction kernel modulation profiles were identified through an alternating optimization algorithm where tube current was updated analytically followed by a gradient-based optimization of reconstruction kernel. The non-circular orbit is first parameterized as a linear combination of bases functions and the coefficients were then optimized using an evolutionary algorithm. The task-driven strategy was compared with conventional acquisitions without modulation, using automatic exposure control, and in a circular orbit. Results: The task-driven strategy outperformed conventional techniques in all tasks investigated, improving the detectability of a spherical lesion detection task by an average of 50% in the interior of a pelvis phantom. The non-circular orbit design successfully mitigated photon starvation effects arising from a dense embolization coil in a head phantom, improving the conspicuity of an intracranial hemorrhage proximal to the coil. Conclusion: The task-driven imaging framework leverages a knowledge of the imaging task within a patient-specific anatomical model to optimize image acquisition and reconstruction techniques, thereby improving imaging performance beyond that achievable with conventional approaches. 2R01-CA-112163; R01-EB-017226; U01-EB-018758; Siemens Healthcare (Forcheim, Germany)« less

Controlling the orbital sequence in individual Cu-phthalocyanine molecules.

PubMed

Uhlmann, C; Swart, I; Repp, J

2013-02-13

We report on the controlled change of the energetic ordering of molecular orbitals. Negatively charged copper(II)phthalocyanine on NaCl/Cu(100) undergoes a Jahn-Teller distortion that lifts the degeneracy of two frontier orbitals. The energetic order of the levels can be controlled by Au and Ag atoms in the vicinity of the molecule. As only one of the states is occupied, the control of the energetic order is accompanied by bistable changes of the charge distribution inside the molecule, rendering it a bistable switch.

KSC-04pd0950

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is lowered toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0948

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is maneuvered toward the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

Analysis of Orbital Lifetime Prediction Parameters in Preparation for Post-Mission Disposal

NASA Astrophysics Data System (ADS)

Choi, Ha-Yeon; Kim, Hae-Dong; Seong, Jae-Dong

2015-12-01

Atmospheric drag force is an important source of perturbation of Low Earth Orbit (LEO) orbit satellites, and solar activity is a major factor for changes in atmospheric density. In particular, the orbital lifetime of a satellite varies with changes in solar activity, so care must be taken in predicting the remaining orbital lifetime during preparation for post-mission disposal. In this paper, the System Tool Kit (STK®) Long-term Orbit Propagator is used to analyze the changes in orbital lifetime predictions with respect to solar activity. In addition, the STK® Lifetime tool is used to analyze the change in orbital lifetime with respect to solar flux data generation, which is needed for the orbital lifetime calculation, and its control on the drag coefficient control. Analysis showed that the application of the most recent solar flux file within the Lifetime tool gives a predicted trend that is closest to the actual orbit. We also examine the effect of the drag coefficient, by performing a comparative analysis between varying and constant coefficients in terms of solar activity intensities.

Automatic Detection and Positioning of Ground Control Points Using TerraSAR-X Multiaspect Acquisitions

NASA Astrophysics Data System (ADS)

Montazeri, Sina; Gisinger, Christoph; Eineder, Michael; Zhu, Xiao xiang

2018-05-01

Geodetic stereo Synthetic Aperture Radar (SAR) is capable of absolute three-dimensional localization of natural Persistent Scatterer (PS)s which allows for Ground Control Point (GCP) generation using only SAR data. The prerequisite for the method to achieve high precision results is the correct detection of common scatterers in SAR images acquired from different viewing geometries. In this contribution, we describe three strategies for automatic detection of identical targets in SAR images of urban areas taken from different orbit tracks. Moreover, a complete work-flow for automatic generation of large number of GCPs using SAR data is presented and its applicability is shown by exploiting TerraSAR-X (TS-X) high resolution spotlight images over the city of Oulu, Finland and a test site in Berlin, Germany.

Overload control of artificial gravity facility using spinning tether system for high eccentricity transfer orbits

NASA Astrophysics Data System (ADS)

Gou, Xing-wang; Li, Ai-jun; Tian, Hao-chang; Wang, Chang-qing; Lu, Hong-shi

2018-06-01

As the major part of space life supporting systems, artificial gravity requires further study before it becomes mature. Spinning tether system is a good alternative solution to provide artificial gravity for the whole spacecraft other than additional devices, and its longer tether length could significantly reduce spinning velocity and thus enhance comfortability. An approximated overload-based feedback method is proposed to provide estimated spinning velocity signals for controller, so that gravity level could be accurately controlled without complicated GPS modules. System behavior in high eccentricity transfer orbits is also studied to give a complete knowledge of the spinning stabilities. The application range of the proposed method is studied in various orbit cases and spinning velocities, indicating that it is accurate and reliable for most of the mission phases especially for the final constant gravity level phase. In order to provide stable gravity level for transfer orbit missions, a sliding mode controller based on estimated angular signals is designed for closed-loop control. Numerical results indicate that the combination of overload-based feedback and sliding mode controller could satisfy most of the long-term artificial gravity missions. It is capable of forming flexible gravity environment in relatively good accuracy even in the lowest possible orbital radiuses and high eccentricity orbits of crewed space missions. The proposed scheme provides an effective tether solution for the artificial gravity construction in interstellar travel.

On Adopting a Proactive Approach to the Disposition of Mars Orbiters

NASA Astrophysics Data System (ADS)

Rummel, John; Shotwell, Robert; Price, Hoppy

2016-07-01

There are currently three U.S. orbiters at Mars plus two orbiters from other nations. At the end of each mission, it is desirable to dispose of the vehicle in a condition where it would present no hazard to other orbiters and to potential future crewed vehicles, while meeting planetary protection constraints. There is currently no way to accurately track and confirm positions of these orbiters after they are no longer being actively tracked from Earth, and due to the extremely "bumpy" nature of the Martian gravity field the position of these vehicles rapidly becomes unknowable . The current COSPAR Planetary Protection Policy for Mars includes a throwback to an earlier era of planetary exploration. The Policy's provisions for the disposition of Mars orbiters includes an option "to meet orbital lifetime requirements" of 20 years (at 99% probability) and 50 years (at 95% probability) after launch (which we will call option 1), or option 2, requiring total (surface, mated, and encapsulated) bioburden levels of ≤ 5 x 10 ^{5} spores - which may be discounted by the number of spores thought to be destroyed during the eventual deorbit of the spacecraft. Reference to DeVincenzi et al (1996) illustrates that the current orbital lifetime requirements in option 1 are directly tied to the notion of a "period of exploration," rather than to any explicit expectation of bioload reduction. Under the current regime, all orbiters comply with option 1, orbital lifetime, or option 2, prior to launch, which generally includes an approved bioburden reduction program prior to launch. As part of option 2, a break up/burn up analysis is also performed, covering the event of an uncontrolled re-entry at arrival or during the orbital mission itself. It has been suggested that we should be seeing an increasing tempo of Mars operations, with an emphasis on making maximum use of communications orbiters in particular. It can be shown that for many orbiters, deorbiting can take quite a bit less delta-V than orbit raising (e.g., to >500 km) to extend their lifetime, so deorbiting (with an attendant focus on increasing burnup and breakup to limit microbial contamination) could provide for possibly years more relay support, as well as a more explicit step for bioburden reduction. Deorbiting can also provide for a more positive termination, eliminates future risk of orbital debris generation around Mars, and ends the possibility of re-entering in an uncontrolled fashion and uncontrolled location later. This paper lays out the issues and options associated with an emphasis on option 2 as the preferred orbiter disposal strategy for Category III missions to Mars. The expectation is that valuable orbits can be better protected for future explorers, and that controls over the contamination of Mars by orbiters can be improved if COSPAR adopts this re-emphasis on behalf of the agencies which are now planning future missions to Mar. Ref. DeVincenzi, D. L., P. D. Stabekis, and J. Barengoltz, Refinement of planetary protection policy for Mars missions, Adv. Space Res. 18: 311-316, 1996.

Compact Micromachined Bandpass Filters for Infrared Planetary Spectroscopy

NASA Technical Reports Server (NTRS)

Brown, Ari D.; Aslam, Shahid; Chervenak, James A.; Huang, Wei-Chung; Merrell, Willie; Quijada, Manuel

2011-01-01

The thermal instrument strawman payload of the Jupiter Europa Orbiter on the Europa Jupiter Science Mission will map out thermal anomalies, the structure, and atmospheric conditions of Europa and Jupiter within the 7-100 micron spectral range. One key requirement for the payload is that the mass cannot exceed 3.7 kg. Consequently, a new generation of light-weight miniaturized spectrometers needs to be developed. On the path toward developing these spectrometers is development of ancillary miniaturized spectroscopic components. In this paper, we present a strategy for making radiation hard and low mass FIR band pass metal mesh filters. Our strategy involves using MEMS-based fabrication techniques, which will permit the quasi-optical filter structures to be made with micron-scale precision. This will enable us to achieve tight control over both the pass band of the filter and the micromachined silicon support structure architecture, which will facilitate integration of the filters for a variety of applications.

Multi-step optimization strategy for fuel-optimal orbital transfer of low-thrust spacecraft

NASA Astrophysics Data System (ADS)

Rasotto, M.; Armellin, R.; Di Lizia, P.

2016-03-01

An effective method for the design of fuel-optimal transfers in two- and three-body dynamics is presented. The optimal control problem is formulated using calculus of variation and primer vector theory. This leads to a multi-point boundary value problem (MPBVP), characterized by complex inner constraints and a discontinuous thrust profile. The first issue is addressed by embedding the MPBVP in a parametric optimization problem, thus allowing a simplification of the set of transversality constraints. The second problem is solved by representing the discontinuous control function by a smooth function depending on a continuation parameter. The resulting trajectory optimization method can deal with different intermediate conditions, and no a priori knowledge of the control structure is required. Test cases in both the two- and three-body dynamics show the capability of the method in solving complex trajectory design problems.

Dimensionality Control of d-orbital Occupation in Oxide Superlattices

PubMed Central

Jeong, Da Woon; Choi, Woo Seok; Okamoto, Satoshi; Kim, Jae–Young; Kim, Kyung Wan; Moon, Soon Jae; Cho, Deok–Yong; Lee, Ho Nyung; Noh, Tae Won

2014-01-01

Manipulating the orbital state in a strongly correlated electron system is of fundamental and technological importance for exploring and developing novel electronic phases. Here, we report an unambiguous demonstration of orbital occupancy control between t2g and eg multiplets in quasi-two-dimensional transition metal oxide superlattices (SLs) composed of a Mott insulator LaCoO3 and a band insulator LaAlO3. As the LaCoO3 sublayer thickness approaches its fundamental limit (i.e. one unit-cell-thick), the electronic state of the SLs changed from a Mott insulator, in which both t2g and eg orbitals are partially filled, to a band insulator by completely filling (emptying) the t2g (eg) orbitals. We found the reduction of dimensionality has a profound effect on the electronic structure evolution, which is, whereas, insensitive to the epitaxial strain. The remarkable orbital controllability shown here offers a promising pathway for novel applications such as catalysis and photovoltaics, where the energy of d level is an essential parameter. PMID:25134975

Chemical control of orbital polarization in artificially structured transition-metal oxides: La2NiXO6 (X=B,Al,Ga,In) from first principles

NASA Astrophysics Data System (ADS)

Han, M. J.; Marianetti, C. A.; Millis, A. J.

2010-10-01

The application of modern layer-by-layer growth techniques to transition-metal oxide materials raises the possibility of creating new classes of materials with rationally designed correlated electron properties. An important step toward this goal is the demonstration that electronic structure can be controlled by atomic composition. In compounds with partially occupied transition-metal d shells, one important aspect of the electronic structure is the relative occupancy of different d orbitals. Previous work has established that strain and quantum confinement can be used to influence orbital occupancy. In this paper we demonstrate a different modality for orbital control in transition-metal oxide heterostructures, using density-functional band calculations supplemented by a tight-binding analysis to show that the choice of nontransition-metal counterion X in transition-metal oxide heterostructures composed of alternating LaNiO3 and LaXO3 units strongly affects orbital occupancy, changing the magnitude and in some cases the sign of the orbital polarization.

Light control of orbital domains: case of the prototypical manganite La0.5Sr1.5MnO4

NASA Astrophysics Data System (ADS)

Miller, Timothy; Gensch, Michael; Wall, Simon

2016-12-01

Control of electronic and structural ordering in correlated materials on the ultrafast timescale with light is a new and emerging approach to disentangle the complex interplay of the charge, spin, orbital and structural degree of freedom. In this paper we present an overview of how orbital order and orbital domains can be controlled by near IR and THz radiation in the layered manganite La0.5Sr1.5MnO4. We show how near-IR pumping can efficiently and rapidly melt orbital ordering. However, the nanoscale domain structure recovers unchanged demonstrating the importance of structural defects for the orbital domain formation. On the contrary, we show that pulsed THz fields can be used to effectively orientate the domains. In this case the alignment depends on the in-plane electric field polarisation and is induced by an energy penalty that arises from THz field induced hopping of the localised charges.

Detail view of the starboard side of the aft fuselage ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of the starboard side of the aft fuselage of the Orbiter Discovery in the Orbiter Processing Facility at Kennedy Space Center with the Orbiter Maneuvering/Reaction Control Systems Pod removed and exposing the insulating foil used to protect the orbiter structure from the heat generated by the maneuvering and reaction control engines. Also note in the view that the aft fuselage access door has bee removed and also note the ground support equipment attached to the T-0 umbilical plate in the lower left of the view. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Graphical techniques to assist in pointing and control studies of orbiting spacecraft

NASA Technical Reports Server (NTRS)

Howell, L. W.; Ruf, J. H.

1986-01-01

Computer generated graphics are developed to assist in the modeling and assessment of pointing and control systems of orbiting spacecraft. Three-dimensional diagrams are constructed of the Earth and of geometrical models which resemble the spacecraft of interest. Orbital positioning of the spacecraft model relative to the Earth and the orbital ground track are then displayed. A star data base is also available which may be used for telescope pointing and star tracker field-of-views to visually assist in spacecraft pointing and control studies. A geometrical model of the Hubble Space Telescope (HST) is constructed and placed in Earth orbit to demonstrate the use of these programs. Simulated star patterns are then displayed corresponding to the primary mirror's FOV and the telescope's star trackers for various telescope orientations with respect to the celestial sphere.

Application of Numerical Integration and Data Fusion in Unit Vector Method

NASA Astrophysics Data System (ADS)

Zhang, J.

2012-01-01

The Unit Vector Method (UVM) is a series of orbit determination methods which are designed by Purple Mountain Observatory (PMO) and have been applied extensively. It gets the conditional equations for different kinds of data by projecting the basic equation to different unit vectors, and it suits for weighted process for different kinds of data. The high-precision data can play a major role in orbit determination, and accuracy of orbit determination is improved obviously. The improved UVM (PUVM2) promoted the UVM from initial orbit determination to orbit improvement, and unified the initial orbit determination and orbit improvement dynamically. The precision and efficiency are improved further. In this thesis, further research work has been done based on the UVM: Firstly, for the improvement of methods and techniques for observation, the types and decision of the observational data are improved substantially, it is also asked to improve the decision of orbit determination. The analytical perturbation can not meet the requirement. So, the numerical integration for calculating the perturbation has been introduced into the UVM. The accuracy of dynamical model suits for the accuracy of the real data, and the condition equations of UVM are modified accordingly. The accuracy of orbit determination is improved further. Secondly, data fusion method has been introduced into the UVM. The convergence mechanism and the defect of weighted strategy have been made clear in original UVM. The problem has been solved in this method, the calculation of approximate state transition matrix is simplified and the weighted strategy has been improved for the data with different dimension and different precision. Results of orbit determination of simulation and real data show that the work of this thesis is effective: (1) After the numerical integration has been introduced into the UVM, the accuracy of orbit determination is improved obviously, and it suits for the high-accuracy data of available observation apparatus. Compare with the classical differential improvement with the numerical integration, its calculation speed is also improved obviously. (2) After data fusion method has been introduced into the UVM, weighted distribution accords rationally with the accuracy of different kinds of data, all data are fully used and the new method is also good at numerical stability and rational weighted distribution.

Guidance and navigation for rendezvous with an uncooperative target

NASA Astrophysics Data System (ADS)

Telaar, J.; Schlaile, C.; Sommer, J.

2018-06-01

This paper presents a guidance strategy for a rendezvous with an uncooperative target. In the applied design reference mission, a spiral approach is commanded ensuring a collision-free relative orbit due to e/i-vector separation. The dimensions of the relative orbit are successively reduced by Δv commands which at the same time improve the observability of the relative state. The navigation is based on line-of-sight measurements. The relative state is estimated by an extended Kalman filter (EKF). The performance of this guidance and navigation strategy is demonstrated by extensive Monte Carlo simulations taking into account all major uncertainties like measurement errors, Δv execution errors, and differential drag.

Highly reusable space transportation: Approaches for reducing ETO launch costs to $100 - $200 per pound of payload

NASA Technical Reports Server (NTRS)

Olds, John R.

1995-01-01

The Commercial Space Transportation Study (CSTS) suggests that considerable market expansion in earth-to-orbit transportation would take place if current launch prices could be reduced to around $400 per pound of payload. If these low prices can be achieved, annual payload delivered to low earth orbit (LEO) is predicted to reach 6.7 million pounds. The primary market growth will occur in communications, government missions, and civil transportation. By establishing a cost target of $100-$200 per pound of payload for a new launch system, the Highly Reusable Space Transportation (HRST) program has clearly set its sights on removing the current restriction on market growth imposed by today's high launch costs. In particular, achieving the goal of $100-$200 per pound of payload will require significant coordinated efforts in (1) marketing strategy development, (2) business planning, (3) system operational strategy, (4) vehicle technical design, and (5) vehicle maintenance strategy.

Human factors requirements for telerobotic command and control: The European Space Agency experimental programme

NASA Technical Reports Server (NTRS)

Stone, Robert J.

1991-01-01

Space Telerobotics research, performed under contract to the European Space Agency (ESA), concerning the execution of human factors experiments, and ultimately leading to the development of a telerobotics test bed, has been carried out since 1985 by a British Consortium consisting of British Aerospace, the United Kingdom Atomic Energy Authority and, more recently, the UK National Advanced Robotics Research Centre. The principal aim of the first study of the series was to derive preliminary requirements for a teleoperation servicing system, with reference to two mission model scenarios. The first scenario introduced the problem of communications time delays, and their likely effect on the ground-based operator in control of a manipulator system on board an unmanned servicing vehicle in Low Earth Orbit. In the second scenario, the operator was located on the NASA Orbiter aft flight deck, supervising the control of a prototype manipulator in the 'servicing' of an experimental payload in the cargo bay area. Human factors analyses centered on defining the requirements for the teleoperator workstation, such as identifying basic ergonomic requirements for workstation and panel layouts, defining teleoperation strategies, developing alphanumeric and graphic screen formats for the supervision or direct control of the manipulator, and the potential applications of expert system technology. The second study for ESA involved an experimental appraisal of some of the important issues highlighted in the first study, for which relevant human factors data did not exist. Of central importance during the second study was the issue of communications time delays and their effect on the manual control of a teleoperated manipulator from a ground-based command and control station.

Orbital Atherectomy for Treating De Novo Severely Calcified Coronary Narrowing (1-Year Results from the Pivotal ORBIT II Trial).

PubMed

Généreux, Philippe; Lee, Arthur C; Kim, Christopher Y; Lee, Michael; Shlofmitz, Richard; Moses, Jeffrey W; Stone, Gregg W; Chambers, Jeff W

2015-06-15

Percutaneous coronary intervention of severely calcified lesions has historically been associated with major adverse cardiac event (MACE) rates as high as 30%. In the ORBIT II (Evaluate the Safety and Efficacy of OAS in Treating Severely Calcified Coronary Lesions) trial, treatment of de novo severely calcified lesions with the Diamondback 360° Coronary Orbital Atherectomy System (OAS) resulted in low rates of procedural and 30-day adverse ischemic events. The long-term results from this trial have not been reported. We sought to determine the 1-year outcomes after orbital atherectomy of severely calcified coronary lesions. ORBIT II was a single-arm trial enrolling 443 subjects at 49 US sites with severely calcified lesions usually excluded from randomized trials. OAS utilizes a centrifugal differential sanding mechanism of action for plaque modification prior to stent implantation. After OAS drug-eluting stents were implanted in 88.2% of the patients. The primary safety end point was 30-day MACE, the composite of cardiac death, myocardial infarction, or target vessel revascularization [TVR]. The present analysis reports the 1-year follow-up results from ORBIT II. One-year data were available in 433 of 443 patients (97.7%), with median follow-up time of 16.7 months. The 1-year MACE rate was 16.4%, including cardiac death (3.0%), myocardial infarction (9.7%), and target vessel revascularization (5.9%). The 1-year target lesion revascularization rate was 4.7%, and stent thrombosis occurred in 1 patient (0.2%). Independent predictors of 1-year MACE and target vessel revascularization were diameter stenosis at baseline and the use of bare-metal stents. In patients with severely calcified lesions who underwent percutaneous coronary intervention, the use of OAS was associated with low rates of 1-year adverse ischemic events compared with historical controls. This finding has important clinical implications for the selection of optimum treatment strategies for patients with severely calcified lesions. Copyright © 2015 The Authors. Published by Elsevier Inc. All rights reserved.

A Modernized Approach to Meet Diversified Earth Observing System (EOS) AM-1 Mission Requirements

NASA Technical Reports Server (NTRS)

Newman, Lauri Kraft; Hametz, Mark E.; Conway, Darrel J.

1998-01-01

From a flight dynamics perspective, the EOS AM-1 mission design and maneuver operations present a number of interesting challenges. The mission design itself is relatively complex for a low Earth mission, requiring a frozen, Sun-synchronous, polar orbit with a repeating ground track. Beyond the need to design an orbit that meets these requirements, the recent focus on low-cost, "lights out" operations has encouraged a shift to more automated ground support. Flight dynamics activities previously performed in special facilities created solely for that purpose and staffed by personnel with years of design experience are now being shifted to the mission operations centers (MOCs) staffed by flight operations team (FOT) operators. These operators' responsibilities include flight dynamics as a small subset of their work; therefore, FOT personnel often do not have the experience to make critical maneuver design decisions. Thus, streamlining the analysis and planning work required for such a complicated orbit design and preparing FOT personnel to take on the routine operation of such a spacecraft both necessitated increasing the automation level of the flight dynamics functionality. The FreeFlyer(trademark) software developed by AI Solutions provides a means to achieve both of these goals. The graphic interface enables users to interactively perform analyses that previously required many parametric studies and much data reduction to achieve the same result. In addition, the fuzzy logic engine .enables the simultaneous evaluation of multiple conflicting constraints, removing the analyst from the loop and allowing the FOT to perform more of the operations without much background in orbit design. Modernized techniques were implemented for EOS AM-1 flight dynamics support in several areas, including launch window determination, orbit maintenance maneuver control strategies, and maneuver design and calibration automation. The benefits of implementing these techniques include increased fuel available for on-orbit maneuvering, a simplified orbit maintenance process to minimize science data downtime, and an automated routine maneuver planning process. This paper provides an examination of the modernized techniques implemented for EOS AM-1 to achieve these benefits.

A modernized approach to meet diversified earth observing system (EOS) AM-1 mission requirements

NASA Technical Reports Server (NTRS)

Newman, Lauri Kraft; Hametz, Mark E.; Conway, Darrel J.

1998-01-01

From a flight dynamics perspective, the EOS AM-1 mission design and maneuver operations present a number of interesting challenges. The mission design itself is relatively complex for a low Earth mission, requiring a frozen, Sun-synchronous, polar orbit with a repeating ground track. Beyond the need to design an orbit that meets these requirements, the recent focus on low-cost, 'lights out' operations has encouraged a shift to more automated ground support. Flight dynamics activities previously performed in special facilities created solely for that purpose and staffed by personnel with years of design experience are now being shifted to the mission operations centers (MOCs) staffed by flight operations team (FOT) operators. These operators' responsibilities include flight dynamics as a small subset of their work; therefore, FOT personnel often do not have the experience to make critical maneuver design decisions. Thus, streamlining the analysis and planning work required for such a complicated orbit design and preparing FOT personnel to take on the routine operation of such a spacecraft both necessitated increasing the automation level of the flight dynamics functionality. The FreeFlyer(TM) software developed by AI Solutions provides a means to achieve both of these goals. The graphic interface enables users to interactively perform analyses that previously required many parametric studies and much data reduction to achieve the same result In addition, the fuzzy logic engine enables the simultaneous evaluation of multiple conflicting constraints, removing the analyst from the loop and allowing the FOT to perform more of the operations without much background in orbit design. Modernized techniques were implemented for EOS AM-1 flight dynamics support in several areas, including launch window determination, orbit maintenance maneuver control strategies, and maneuver design and calibration automation. The benefits of implementing these techniques include increased fuel available for on-orbit maneuvering, a simplified orbit maintenance process to minimize science data downtime, and an automated routine maneuver planning process. This paper provides an examination of the modernized techniques implemented for EOS AM-1 to achieve these benefits.

Transgene expression patterns indicate that spaceflight affects stress signal perception and transduction in arabidopsis.

PubMed

Paul, A L; Daugherty, C J; Bihn, E A; Chapman, D K; Norwood, K L; Ferl, R J

2001-06-01

The use of plants as integral components of life support systems remains a cornerstone of strategies for long-term human habitation of space and extraterrestrial colonization. Spaceflight experiments over the past few decades have refined the hardware required to grow plants in low-earth orbit and have illuminated fundamental issues regarding spaceflight effects on plant growth and development. Potential incipient hypoxia, resulting from the lack of convection-driven gas movement, has emerged as a possible major impact of microgravity. We developed transgenic Arabidopsis containing the alcohol dehydrogenase (Adh) gene promoter linked to the beta-glucuronidase (GUS) reporter gene to address specifically the possibility that spaceflight induces the plant hypoxia response and to assess whether any spaceflight response was similar to control terrestrial hypoxia-induced gene expression patterns. The staining patterns resulting from a 5-d mission on the orbiter Columbia during mission STS-93 indicate that the Adh/GUS reporter gene was activated in roots during the flight. However, the patterns of expression were not identical to terrestrial control inductions. Moreover, although terrestrial hypoxia induces Adh/GUS expression in the shoot apex, no apex staining was observed in the spaceflight plants. This indicates that either the normal hypoxia response signaling is impaired in spaceflight or that spaceflight inappropriately induces Adh/GUS activity for reasons other than hypoxia.

Transgene expression patterns indicate that spaceflight affects stress signal perception and transduction in arabidopsis

NASA Technical Reports Server (NTRS)

Paul, A. L.; Daugherty, C. J.; Bihn, E. A.; Chapman, D. K.; Norwood, K. L.; Ferl, R. J.

2001-01-01

The use of plants as integral components of life support systems remains a cornerstone of strategies for long-term human habitation of space and extraterrestrial colonization. Spaceflight experiments over the past few decades have refined the hardware required to grow plants in low-earth orbit and have illuminated fundamental issues regarding spaceflight effects on plant growth and development. Potential incipient hypoxia, resulting from the lack of convection-driven gas movement, has emerged as a possible major impact of microgravity. We developed transgenic Arabidopsis containing the alcohol dehydrogenase (Adh) gene promoter linked to the beta-glucuronidase (GUS) reporter gene to address specifically the possibility that spaceflight induces the plant hypoxia response and to assess whether any spaceflight response was similar to control terrestrial hypoxia-induced gene expression patterns. The staining patterns resulting from a 5-d mission on the orbiter Columbia during mission STS-93 indicate that the Adh/GUS reporter gene was activated in roots during the flight. However, the patterns of expression were not identical to terrestrial control inductions. Moreover, although terrestrial hypoxia induces Adh/GUS expression in the shoot apex, no apex staining was observed in the spaceflight plants. This indicates that either the normal hypoxia response signaling is impaired in spaceflight or that spaceflight inappropriately induces Adh/GUS activity for reasons other than hypoxia.

Radio interferometric measurements for accurate planetary orbiter navigation

NASA Technical Reports Server (NTRS)

Poole, S. R.; Ananda, M.; Hildebrand, C. E.

1979-01-01

The use of narrowband delta-VLBI to achieve accurate orbit determination is presented by viewing a spacecraft from widely separated stations followed by viewing a nearby quasar from the same stations. Current analysis is examined that establishes the orbit determination accuracy achieved with data arcs spanning up to 3.5 d. Strategies for improving prediction accuracy are given, and the performance of delta-VLBI is compared with conventional radiometric tracking data. It is found that accuracy 'within the fit' is on the order of 0.5 km for data arcs having delta-VLBI on the ends of the arcs and for arc lengths varying from one baseline to 3.5 d. The technique is discussed with reference to the proposed Venus Orbiting Imaging Radar mission.

Orbiter integrated active thermal control subsystem test

NASA Technical Reports Server (NTRS)

Jaax, J. R.

1980-01-01

Integrated subsystem level testing of the systems within the orbiter active thermal chamber capable of simulating ground, orbital, and entry temperature and pressure profiles. The test article was in a closed loop configuration that included flight type and functionally simulated protions of all ATCS components for collecting, transporting, and rejecting orbiter waste heat. Specially designed independently operating equipment simulated the transient thermal input from the cabin, payload, fuel cells, freon cold plates, hydraulic system, and space environment. Test team members using data, controls, and procedures available to a flight crew controlled the operation of the ATCS. The ATCS performance met or exceeded all thermal and operational requirements for planned and contingency mission support.

Lay out, test verification and in orbit performance of HELIOS a temperature control system

NASA Technical Reports Server (NTRS)

Brungs, W.

1975-01-01

HELIOS temperature control system is described. The main design features and the impact of interactions between experiment, spacecraft system, and temperature control system requirements on the design are discussed. The major limitations of the thermal design regarding a closer sun approach are given and related to test experience and performance data obtained in orbit. Finally the validity of the test results achieved with prototype and flight spacecraft is evaluated by comparison between test data, orbit temperature predictions and flight data.

An autonomous payload controller for the Space Shuttle

NASA Technical Reports Server (NTRS)

Hudgins, J. I.

1979-01-01

The Autonomous Payload Control (APC) system discussed in the present paper was designed on the basis of such criteria as minimal cost of implementation, minimal space required in the flight-deck area, simple operation with verification of the results, minimal additional weight, minimal impact on Orbiter design, and minimal impact on Orbiter payload integration. In its present configuration, the APC provides a means for the Orbiter crew to control as many as 31 autononous payloads. The avionics and human engineering aspects of the system are discussed.

Cost-Effective Telemetry and Command Ground Systems Automation Strategy for the Soil Moisture Active Passive (SMAP) Mission

NASA Technical Reports Server (NTRS)

Choi, Josh; Sanders, Antonio

2012-01-01

Soil Moisture Active Passive (SMAP) is an Earth-orbiting, remote-sensing NASA mission slated for launch in 2014. The ground data system (GDS) being developed for SMAP is composed of many heterogeneous subsystems, ranging from those that support planning and sequencing to those used for real-time operations, and even further to those that enable science data exchange. A full end-to-end automation of the GDS may result in cost savings during mission operations, but it would require a significant upfront investment to develop such a comprehensive automation. As demonstrated by the Jason-1 and Wide-field Infrared Survey Explorer (WISE) missions, a measure of "lights-out" automation for routine, orbital pass, ground operations can still reduce mission costs through smaller staffing of operators and limiting their working hours. The challenge, then, for the SMAP GDS engineering team, is to formulate an automated operations strategy--and corresponding system architecture -- to minimize operator intervention during routine operations, while balancing the development costs associated with the scope and complexity of automation. This paper discusses the automated operations approach being developed for the SMAP GDS. The focus is on automating the activities involved in routine passes, which limits the scope to real-time operations. A key subsystem of the SMAP GDS -- NASA's AMMOS Mission Data Processing and Control System (AMPCS) -- provides a set of capabilities that enable such automation. Also discussed are the lights-out pass automations of the Jason-1 and WISE missions and how they informed the automation strategy for SMAP. The paper aims to provide insights into what is necessary in automating the GDS operations for Earth satellite missions.

Cost-Effective Telemetry and Command Ground Systems Automation Strategy for the Soil Moisture Active Passive (SMAP) Mission

NASA Technical Reports Server (NTRS)

Choi, Joshua S.; Sanders, Antonio L.

2012-01-01

Soil Moisture Active Passive (SMAP) is an Earth-orbiting, remote-sensing NASA mission slated for launch in 2014.[double dagger] The ground data system (GDS) being developed for SMAP is composed of many heterogeneous subsystems, ranging from those that support planning and sequencing to those used for real-time operations, and even further to those that enable science data exchange. A full end-to-end automation of the GDS may result in cost savings during mission operations, but it would require a significant upfront investment to develop such comprehensive automation. As demonstrated by the Jason-1 and Wide-field Infrared Survey Explorer (WISE) missions, a measure of "lights-out" automation for routine, orbital pass ground operations can still reduce mission cost through smaller staffing of operators and limited work hours. The challenge, then, for the SMAP GDS engineering team is to formulate an automated operations strategy--and corresponding system architecture--to minimize operator intervention during operations, while balancing the development cost associated with the scope and complexity of automation. This paper discusses the automated operations approach being developed for the SMAP GDS. The focus is on automating the activities involved in routine passes, which limits the scope to real-time operations. A key subsystem of the SMAP GDS--NASA's AMMOS Mission Data Processing and Control System (AMPCS)--provides a set of capabilities that enable such automation. Also discussed are the lights-out pass automations of the Jason-1 and WISE missions and how they informed the automation strategy for SMAP. The paper aims to provide insights into what is necessary in automating the GDS operations for Earth satellite missions.

Semiautomatic regional segmentation to measure orbital fat volumes in thyroid-associated ophthalmopathy. A validation study.

PubMed

Comerci, M; Elefante, A; Strianese, D; Senese, R; Bonavolontà, P; Alfano, B; Bonavolontà, B; Brunetti, A

2013-08-01

This study was designed to validate a novel semi-automated segmentation method to measure regional intra-orbital fat tissue volume in Graves' ophthalmopathy. Twenty-four orbits from 12 patients with Graves' ophthalmopathy, 24 orbits from 12 controls, ten orbits from five MRI study simulations and two orbits from a digital model were used. Following manual region of interest definition of the orbital volumes performed by two operators with different levels of expertise, an automated procedure calculated intra-orbital fat tissue volumes (global and regional, with automated definition of four quadrants). In patients with Graves' disease, clinical activity score and degree of exophthalmos were measured and correlated with intra-orbital fat volumes. Operator performance was evaluated and statistical analysis of the measurements was performed. Accurate intra-orbital fat volume measurements were obtained with coefficients of variation below 5%. The mean operator difference in total fat volume measurements was 0.56%. Patients had significantly higher intra-orbital fat volumes than controls (p<0.001 using Student's t test). Fat volumes and clinical score were significantly correlated (p<0.001). The semi-automated method described here can provide accurate, reproducible intra-orbital fat measurements with low inter-operator variation and good correlation with clinical data.

Control of the orbital ordering in manganite superlattices and impact on properties

NASA Astrophysics Data System (ADS)

Koçak, Ayşegül Begüm; Varignon, Julien; Lemal, Sébastien; Ghosez, Philippe; Lepetit, Marie-Bernadette

2017-09-01

The present paper theoretically studies the possibility to control the orbital ordering in manganite superlattices. Indeed, favored dz2eg -orbital occupancy is one of the proposed interpretations for the formation of a "dead" layer at the interfaces in manganite thin films and superlattices. We show here that favored dz2eg -orbital occupancy at the interfaces can be prevented by using alkaline-earth simple oxides as alternating layers in very thin superlattices. Such an alternating layer promotes the contraction of the manganite layers at the interfaces and favors a dx2-y2eg orbital occupancy. This result holds for different manganites, different alkaline-earth simple oxides, as well as different thicknesses of the two layers. While Boltzmann's transport calculations on different superlattices show unexpectedly only weak dependence of the electrical conductivity on the orbital ordering, the enhanced occupation of the dx2-y2 orbital should result in an increased Curie temperature.

Slow Orbit Feedback at the ALS Using Matlab

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

Portmann, G.

1999-03-25

The third generation Advanced Light Source (ALS) produces extremely bright and finely focused photon beams using undulatory, wigglers, and bend magnets. In order to position the photon beams accurately, a slow global orbit feedback system has been developed. The dominant causes of orbit motion at the ALS are temperature variation and insertion device motion. This type of motion can be removed using slow global orbit feedback with a data rate of a few Hertz. The remaining orbit motion in the ALS is only 1-3 micron rms. Slow orbit feedback does not require high computational throughput. At the ALS, the globalmore » orbit feedback algorithm, based on the singular valued decomposition method, is coded in MATLAB and runs on a control room workstation. Using the MATLAB environment to develop, test, and run the storage ring control algorithms has proven to be a fast and efficient way to operate the ALS.« less

Impingement effect of service module reaction control system engine plumes. Results of service module reaction control system plume model force field application to an inflight Skylab mission proximity operation situation with the inflight Skylab response

NASA Technical Reports Server (NTRS)

Lobb, J. D., Jr.

1978-01-01

Plume impingement effects of the service module reaction control system thruster firings were studied to determine if previous flight experience would support the current plume impingement model for the orbiter reaction control system engines. The orbiter reaction control system is used for rotational and translational maneuvers such as those required during rendezvous, braking, docking, and station keeping. Therefore, an understanding of the characteristics and effects of the plume force fields generated by the reaction control system thruster firings were examined to develop the procedures for orbiter/payload proximity operations.

Application of GPS tracking techniques to orbit determination for TDRS

NASA Technical Reports Server (NTRS)

Haines, B. J.; Lichten, S. M.; Malla, R. P.; Wu, S. C.

1993-01-01

In this paper, we evaluate two fundamentally different approaches to TDRS orbit determination utilizing Global Positioning System (GPS) technology and GPS-related techniques. In the first, a GPS flight receiver is deployed on the TDRSS spacecraft. The TDRS ephemerides are determined using direct ranging to the GPS spacecraft, and no ground network is required. In the second approach, the TDRSS spacecraft broadcast a suitable beacon signal, permitting the simultaneous tracking of GPS and TDRSS satellites from a small ground network. Both strategies can be designed to meet future operational requirements for TDRS-2 orbit determination.

A review of GPS-based tracking techniques for TDRS orbit determination

NASA Technical Reports Server (NTRS)

Haines, B. J.; Lichten, S. M.; Malla, R. P.; Wu, S.-C.

1993-01-01

This article evaluates two fundamentally different approaches to the Tracking and Data Relay Satellite (TDRS) orbit determination utilizing Global Positioning System (GPS) technology and GPS-related techniques. In the first, a GPS flight receiver is deployed on the TDRS. The TDRS ephemerides are determined using direct ranging to the GPS spacecraft, and no ground network is required. In the second approach, the TDRS's broadcast a suitable beacon signal, permitting the simultaneous tracking of GPS and Tracking and Data Relay Satellite System satellites by ground receivers. Both strategies can be designed to meet future operational requirements for TDRS-II orbit determination.

Optimizing Spacecraft Placement for Liaison Constellations

NASA Technical Reports Server (NTRS)

Chow, C. Channing; Villac, Benjamin F.; Lo, Martin W.

2011-01-01

A navigation and communications network is proposed to support an anticipated need for infrastructure in the Earth-Moon system. Periodic orbits will host the constellations while a novel, autonomous navigation strategy will guide the spacecraft along their path strictly based on satellite-to-satellite telemetry. In particular, this paper investigates the second stage of a larger constellation optimization scheme for multi-spacecraft systems. That is, following an initial orbit down-selection process, this analysis provides insights into the ancillary problem of spacecraft placement. Two case studies are presented that consider configurations of up to four spacecraft for a halo orbit and a cycler trajectory.

Ballistics, Navigation and Motion Control of the S/C on Stages of the Phobos Surface Approaching and Landing

NASA Astrophysics Data System (ADS)

Akim, E. L.; Ruzskiy, E. G.; Shishov, V. A.; Stepaniants, V. A.; Tuchin, A. G.

The main purpose of the Federal space program «Phobos-grunt» project is to deliver samples of a ground of Mars's natural satellite Phobos on Earth. Mission is planned to begin at the start window of 2009 year and to finish at 2011 year. The work is devoted to problems connected with development of the Phobos landing scheme. The major factors, the landing strategy is formed on, are the tasks connected with orbital mechanics, restrictions obliged by the space craft (SC) onboard systems, inexact knowledge of the Phobos's kinematics parameters, its hypsometry and gravity field.

Earth observing system - Concepts and implementation strategy

NASA Technical Reports Server (NTRS)

Hartle, R. E.

1986-01-01

The concepts of an Earth Observing System (EOS), an information system being developed by the EOS Science and Mission Requirements Working Group for international use and planned to begin in the 1990s, are discussed. The EOS is designed to study the factors that control the earth's hydrologic cycle, biochemical cycles, and climatologic processes by combining the measurements from remote sensing instruments, in situ measurement devices, and a data and information system. Three EOS platforms are planned to be launched into low, polar, sun-synchronous orbits during the Space Station's Initial Operating Configuration, one to be provided by ESA and two by the United States.

Simple control laws for low-thrust orbit transfers

NASA Technical Reports Server (NTRS)

Petropoulos, Anastassios E.

2003-01-01

Two methods are presented by which to determine both a thrust direction and when to apply thrust to effect specified changes in any of the orbit elements except for true anomaly, which is assumed free. The central body is assumed to be a point mass, and the initial and final orbits are assumed closed. Thrust, when on, is of a constant value, and specific impulse is constant. The thrust profiles derived from the two methods are not propellant-optimal, but are based firstly on the optimal thrust directions and location on the osculating orbit for changing each of the orbit elements and secondly on the desired changes in the orbit elements. Two examples of transfers are presented, one in semimajor axis and inclination, and one in semimajor axis and eccentricity. The latter compares favourably with a propellant-optimized transfer between the same orbits. The control laws have few input parameters, but can still capture the complexity of a wide variety of orbit transfers.

Guidance Scheme for Modulation of Drag Devices to Enable Return from Low Earth Orbit

NASA Technical Reports Server (NTRS)

Dutta, Soumyo; Bowes, Angela L.; Cianciolo, Alicia D.; Glass, Christopher E.; Powell, Richard W.

2017-01-01

Passive drag devices provide opportunities to return payloads from low Earth orbits quickly without using onboard propulsive systems to de-orbit the spacecraft. However, one potential disadvantage of such systems has been the lack of landing accuracy. Drag modulation or changing the shape of the drag device during flight offer a way to control the de-orbit trajectory and target a specific landing location. This paper discusses a candidate passive drag based system, called Exo-brake, as well as efforts to model the dynamics of the vehicle as it de-orbits and guidance schemes used to control the trajectory. Such systems can enable quick return of payloads from low Earth orbit assets like the International Space Station without the use of large re-entry cargo capsules or propulsive systems.

Fuzzy attitude control for a nanosatellite in leo orbit

NASA Astrophysics Data System (ADS)

Calvo, Daniel; Laverón-Simavilla, Ana; Lapuerta, Victoria; Aviles, Taisir

Fuzzy logic controllers are flexible and simple, suitable for small satellites Attitude Determination and Control Subsystems (ADCS). In this work, a tailored fuzzy controller is designed for a nanosatellite and is compared with a traditional Proportional Integrative Derivative (PID) controller. Both control methodologies are compared within the same specific mission. The orbit height varies along the mission from injection at around 380 km down to a 200 km height orbit, and the mission requires pointing accuracy over the whole time. Due to both the requirements imposed by such a low orbit, and the limitations in the power available for the attitude control, a robust and efficient ADCS is required. For these reasons a fuzzy logic controller is implemented as the brain of the ADCS and its performance and efficiency are compared to a traditional PID. The fuzzy controller is designed in three separated controllers, each one acting on one of the Euler angles of the satellite in an orbital frame. The fuzzy memberships are constructed taking into account the mission requirements, the physical properties of the satellite and the expected performances. Both methodologies, fuzzy and PID, are fine-tuned using an automated procedure to grant maximum efficiency with fixed performances. Finally both methods are probed in different environments to test their characteristics. The simulations show that the fuzzy controller is much more efficient (up to 65% less power required) in single maneuvers, achieving similar, or even better, precision than the PID. The accuracy and efficiency improvement of the fuzzy controller increase with orbit height because the environmental disturbances decrease, approaching the ideal scenario. A brief mission description is depicted as well as the design process of both ADCS controllers. Finally the validation process and the results obtained during the simulations are described. Those results show that the fuzzy logic methodology is valid for small satellites' missions benefiting from a well-developed artificial intelligence theory.

Characterizing the Survey Strategy and Initial Orbit Determination Abilities of the NASA MCAT Telescope for Geosynchronous Orbital Debris Environmental Studies

NASA Technical Reports Server (NTRS)

Frith, James; Barker, Ed; Cowardin, Heather; Buckalew, Brent; Anz-Meado, Phillip; Lederer, Susan

2017-01-01

The NASA Orbital Debris Program Office (ODPO) recently commissioned the Meter Class Autonomous Telescope (MCAT) on Ascension Island with the primary goal of obtaining population statistics of the geosynchronous (GEO) orbital debris environment. To help facilitate this, studies have been conducted using MCAT's known and projected capabilities to estimate the accuracy and timeliness in which it can survey the GEO environment. A simulated GEO debris population is created and sampled at various cadences and run through the Constrained Admissible Region Multi Hypotheses Filter (CAR-MHF). The orbits computed from the results are then compared to the simulated data to assess MCAT's ability to determine accurately the orbits of debris at various sample rates. Additionally, estimates of the rate at which MCAT will be able produce a complete GEO survey are presented using collected weather data and the proposed observation data collection cadence. The specific methods and results are presented here.

The role of predicted solar activity in TOPEX/Poseidon orbit maintenance maneuver design

NASA Technical Reports Server (NTRS)

Frauenholz, Raymond B.; Shapiro, Bruce E.

1992-01-01

Following launch in June 1992, the TOPEX/Poseidon satellite will be placed in a near-circular frozen orbit at an altitude of about 1336 km. Orbit maintenance maneuvers are planned to assure all nodes of the 127-orbit 10-day repeat ground track remain within a 2 km equatorial longitude bandwidth. Orbit determination, maneuver execution, and atmospheric drag prediction errors limit overall targeting performance. This paper focuses on the effects of drag modeling errors, with primary emphasis on the role of SESC solar activity predictions, especially the 27-day outlook of the 10.7 cm solar flux and geomagnetic index used by a simplified version of the Jacchia-Roberts density model developed for this TOPEX/Poseidon application. For data evaluated from 1983-90, the SESC outlook performed better than a simpler persistence strategy, especially during the first 7-10 days. A targeting example illustrates the use of ground track biasing to compensate for expected orbit predictions errors, emphasizing the role of solar activity prediction errors.

Precise regional baseline estimation using a priori orbital information

NASA Technical Reports Server (NTRS)

Lindqwister, Ulf J.; Lichten, Stephen M.; Blewitt, Geoffrey

1990-01-01

A solution using GPS measurements acquired during the CASA Uno campaign has resulted in 3-4 mm horizontal daily baseline repeatability and 13 mm vertical repeatability for a 729 km baseline, located in North America. The agreement with VLBI is at the level of 10-20 mm for all components. The results were obtained with the GIPSY orbit determination and baseline estimation software and are based on five single-day data arcs spanning the 20, 21, 25, 26, and 27 of January, 1988. The estimation strategy included resolving the carrier phase integer ambiguities, utilizing an optial set of fixed reference stations, and constraining GPS orbit parameters by applying a priori information. A multiday GPS orbit and baseline solution has yielded similar 2-4 mm horizontal daily repeatabilities for the same baseline, consistent with the constrained single-day arc solutions. The application of weak constraints to the orbital state for single-day data arcs produces solutions which approach the precise orbits obtained with unconstrained multiday arc solutions.

A case report of orbital Langerhans cell histiocytosis presenting as a orbital cellulitis.

PubMed

Albert-Fort, M; González-Candial, M

2018-04-08

A 10-year-old girl was seen with a 3-week history of right upper lid swelling and with no other symptoms or fever. There was no recent history of sinusitis, trauma, or previous infection involving the periorbital area, or response to oral antibiotic treatment. Orbital computed tomography showed a lesion involving the upper margin of the orbit, and bone destruction at the orbital roof. Biopsy performed revealed the presence of Langerhans cell Histiocytosis. The lesion was surgically debulked and corticosteroids were used intra-operatively. The lesion responded to treatment. The orbital involvement of Langerhans cell histiocytosis, despite its low incidence, should be considered in the examination of acute peri-orbital swelling. It usually presents as an osteolytic lesion, and it is confirmed with a histological examination and immunohistochemical techniques for CD1a and S100. An interdisciplinary approach is recommended to rule out multifocal or multisystemic diseases, as well as to develop an appropriate treatment strategy. Copyright © 2018 Sociedad Española de Oftalmología. Publicado por Elsevier España, S.L.U. All rights reserved.

KSC-04pd0942

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod (seen from the back) is lifted off its transporter. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0949

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers stand by as the left-hand Orbital Maneuvering System (OMS) pod is maneuvered toward the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0944

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is lifted at an angle from the transporter below. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0947

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod (top of photo) is poised behind the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0946

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers on an upper level watch as the left-hand Orbital Maneuvering System (OMS) pod is lifted high to maneuver it toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0943

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers check the lifting of the left-hand Orbital Maneuvering System (OMS) pod. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0941

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod (seen from the front) is lifted off its transporter. The OMS pod will be installed on the orbiter Discovery. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0945

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, a worker on an upper level watches as the left-hand Orbital Maneuvering System (OMS) pod is lifted high to maneuver it toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

STS-119 Flight Control Team in WFCR - Orbit 3 - Flight Director Bryan Lunney

NASA Image and Video Library

2009-03-24

JSC2009-E-061542 (24 March 2009) --- The members of the STS-119 Orbit 3 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA?s Johnson Space Center. Flight director Bryan Lunney (center) near the front.

STS-124/1J ISS Orbit 3 flight control team portrait

NASA Image and Video Library

2008-06-09

JSC2008-E-045777 (9 June 2008) --- The members of the STS-124/1J ISS Orbit 3 flight control team pose for a group portrait in the space station flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Emily Nelson stands in the center foreground.

STS-125 Flight Control Team in WFCR - Orbit 1 - Flight Director Tony Ceccacci

NASA Image and Video Library

2009-05-20

JSC2009-E-120813 (20 May 2009) --- The members of the STS-125 Orbit 1 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Tony Ceccacci holds the STS-125 mission logo.

STS-132 Flight Control Team in WFCR - Orbit 1

NASA Image and Video Library

2010-05-22

JSC2010-E-086698 (22 May 2010) --- The members of the STS-132 Orbit 1 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Mike Sarafin (center) is visible on the front row.

STS-131 Flight Control Team in WFCR - Orbit 2 - Flight Director Mike Sarafin

NASA Image and Video Library

2010-04-14

JSC2010-E-051978 (14 April 2010) --- The members of the STS-131 Orbit 2 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Mike Sarafin holds the STS-131 mission logo.

STS-132 Flight Control Team in WFCR - Orbit 2

NASA Image and Video Library

2010-05-20

JSC2010-E-086451 (20 May 2010) --- The members of the STS-132 Orbit 2 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Chris Edelen (second left) is visible on the front row.

A non-ideal portal frame energy harvester controlled using a pendulum

NASA Astrophysics Data System (ADS)

Iliuk, I.; Balthazar, J. M.; Tusset, A. M.; Piqueira, J. R. C.; Rodrigues de Pontes, B.; Felix, J. L. P.; Bueno, Á. M.

2013-09-01

A model of energy harvester based on a simple portal frame structure is presented. The system is considered to be non-ideal system (NIS) due to interaction with the energy source, a DC motor with limited power supply and the system structure. The nonlinearities present in the piezoelectric material are considered in the piezoelectric coupling mathematical model. The system is a bi-stable Duffing oscillator presenting a chaotic behavior. Analyzing the average power variation, and bifurcation diagrams, the value of the control variable that optimizes power or average value that stabilizes the chaotic system in the periodic orbit is determined. The control sensitivity is determined to parametric errors in the damping and stiffness parameters of the portal frame. The proposed passive control technique uses a simple pendulum to tuned to the vibration of the structure to improve the energy harvesting. The results show that with the implementation of the control strategy it is possible to eliminate the need for active or semi active control, usually more complex. The control also provides a way to regulate the energy captured to a desired operating frequency.

Integrated Digital Flight Control System for the Space Shuttle Orbiter

NASA Technical Reports Server (NTRS)

1973-01-01

The objectives of the integrated digital flight control system (DFCS) is to provide rotational and translational control of the space shuttle orbiter in all phases of flight: from launch ascent through orbit to entry and touchdown, and during powered horizontal flights. The program provides a versatile control system structure while maintaining uniform communications with other programs, sensors, and control effectors by using an executive routine/functional subroutine format. The program reads all external variables at a single point, copies them into its dedicated storage, and then calls the required subroutines in the proper sequence. As a result, the flight control program is largely independent of other programs in the computer complex and is equally insensitive to characteristics of the processor configuration. The integrated structure is described of the control system and the DFCS executive routine which embodies that structure. The input and output, including jet selection are included. Specific estimation and control algorithm are shown for the various mission phases: cruise (including horizontal powered flight), entry, on-orbit, and boost. Attitude maneuver routines that interface with the DFCS are included.

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

NASA Astrophysics Data System (ADS)

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

2015-02-01

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

STS-106 Orbit 1 Flight Team with Phil Engelhauf in WFCR

NASA Image and Video Library

2000-09-15

JSC2000-06253 (15 Sept. 2000) --- Flight director Phil Engelauf, front center, and the other fifty-odd flight controllers making up the STS-106 Orbit 1 team, pose for their group portrait in the Flight Control Room of Houston's Mission Control Center.

Characterizing the Survey Strategy and Initial Orbit Determination Abilities of the NASA MCAT Telescope for Geosynchronous Orbital Debris Environmental Studies

NASA Astrophysics Data System (ADS)

Frith, J.; Barker, E.; Cowardin, H.; Buckalew, B.; Anz-Meador, P.; Lederer, S.

The National Aeronautics and Space Administration (NASA) Orbital Debris Program Office (ODPO) recently commissioned the Meter Class Autonomous Telescope (MCAT) on Ascension Island with the primary goal of obtaining population statistics of the geosynchronous (GEO) orbital debris environment. To help facilitate this, studies have been conducted using MCAT’s known and projected capabilities to estimate the accuracy and timeliness in which it can survey the GEO environment, including collected weather data and the proposed observational data collection cadence. To optimize observing cadences and probability of detection, on-going work using a simulated GEO debris population sampled at various cadences are run through the Constrained Admissible Region Multi Hypotheses Filter (CAR-MHF). The orbits computed from the results are then compared to the simulated data to assess MCAT’s ability to determine accurately the orbits of debris at various sample rates. The goal of this work is to discriminate GEO and near-GEO objects from GEO transfer orbit objects that can appear as GEO objects in the environmental models due to the short arc observation and an assumed circular orbit. The specific methods and results are presented here.

A Survey Of Earth-Moon Libration Orbits: Stationkeeping Strategies And Intra-Orbit Transfers

NASA Technical Reports Server (NTRS)

Folta, David; Vaughn, Frank

2004-01-01

Cislunar space is a readily accessible region that may well develop into a prime staging area in the effort to colonize space near Earth or to colonize the Moon. While there have been statements made by various NASA programs regarding placement of resources in orbit about the Earth-Moon Lagrangian locations, there is no survey of the total cost associated with attaining and maintaining these unique orbits in an operational fashion. Transfer trajectories between these orbits required for assembly, servicing, and positioning of these resources have not been extensively investigated. These orbits are dynamically similar to those used for the Sun-Earth missions, but differences in governing gravitational ratios and perturbation sources result in unique characteristics. We implement numerical computations using high fidelity models and linear and nonlinear targeting techniques to compute the various maneuver (Delta)V and temporal costs associated with orbits about each of the Earth-Moon Lagrangian locations (L1, L2, L3, L4, and L5). From a dynamical system standpoint, we speak to the nature of these orbits and their stability. We address the cost of transfers between each pair of Lagrangian locations.

Lissajous Orbit Control for the Deep Space Climate Observatory Sun-Earth L1 Libration Point Mission

NASA Technical Reports Server (NTRS)

Roberts, Craig; Case, Sarah; Reagoso, John

2015-01-01

DSCOVR Lissajous Orbit sized such that orbit track never extends beyond 15 degrees from Earth-Sun line (as seen from Earth). Requiring delta-V maneuvers, control orbit to obey a Solar Exclusion Zone (SEZ) cone of half-angle 4 degrees about the Earth-Sun line. Spacecraft should never be less than 4 degrees from solar center as seen from Earth. Following Lissajous Orbit Insertion (LOI), DSCOVR should be in an opening phase that just skirts the 4-degree SEZ. Maximizes time to the point where a closing Lissajous will require avoidance maneuvers to keep it out of the SEZ. Station keeping maneuvers should take no more than 15 minutes.

Formation Flying for Distributed InSAR

NASA Technical Reports Server (NTRS)

Scharf, Daniel P.; Murray, Emmanuell A.; Ploen, Scott R.; Gromov, Konstantin G.; Chen, Curtis W.

2006-01-01

We consider two spacecraft flying in formation to create interferometric synthetic aperture radar (InSAR). Several candidate orbits for such in InSar formation have been previously determined based on radar performance and Keplerian orbital dynamics. However, with out active control, disturbance-induced drift can degrade radar performance and (in the worst case) cause a collision. This study evaluates the feasibility of operating the InSAR spacecraft as a formation, that is, with inner-spacecraft sensing and control. We describe the candidate InSAR orbits, design formation guidance and control architectures and algorithms, and report the (Delta)(nu) and control acceleration requirements for the candidate orbits for several tracking performance levels. As part of determining formation requirements, a formation guidance algorithm called Command Virtual Structure is introduced that can reduce the (Delta)(nu) requirements compared to standard Leader/Follower formation approaches.

Controlling the superconducting transition by spin-orbit coupling

NASA Astrophysics Data System (ADS)

Banerjee, N.; Ouassou, J. A.; Zhu, Y.; Stelmashenko, N. A.; Linder, J.; Blamire, M. G.

2018-05-01

Whereas considerable evidence exists for the conversion of singlet Cooper pairs into triplet Cooper pairs in the presence of inhomogeneous magnetic fields, recent theoretical proposals have suggested an alternative way to exert control over triplet generation: intrinsic spin-orbit coupling in a homogeneous ferromagnet coupled to a superconductor. Here, we proximity couple Nb to an asymmetric Pt/Co/Pt trilayer, which acts as an effective spin-orbit-coupled ferromagnet owing to structural inversion asymmetry. Unconventional modulation of the superconducting critical temperature as a function of in-plane and out-of-plane applied magnetic fields suggests the presence of triplets that can be controlled by the magnetic orientation of a single homogeneous ferromagnet. Our studies demonstrate an active role of spin-orbit coupling in controlling the triplets, an important step towards the realization of novel superconducting spintronic devices.

47 CFR 25.280 - Inclined orbit operations.

Code of Federal Regulations, 2010 CFR

2010-10-01

... of the Earth and centered on the satellite's designated service area; (2) Control all electrical... inclined orbit; and (4) Continue to maintain the space station at the authorized longitude orbital location...

Reducing the Risk of Human Missions to Mars Through Testing

NASA Astrophysics Data System (ADS)

Drake, Bret G.

2007-07-01

During the summer of 2002 the NASA Deputy Administrator charted an internal NASA planning group to develop the rationale for exploration beyond low-Earth orbit. This team, termed the Exploration Blueprint, performed architecture analyses to develop roadmaps for how to accomplish the first steps beyond Low-Earth Orbit through the human exploration of Mars. The previous NASA Exploration Team (NEXT) activities laid the foundation and framework for development of NASA s Integrated Space Plan. The reference missions resulting from the analysis performed by the Exploration Blueprint team formed the basis for requirement definition, systems development, technology roadmapping, and risk assessments for future human exploration beyond low-Earth orbit. Emphasis was placed on developing recommendations on what could be done now to effect future exploration activities. The Exploration Blueprint team embraced the Stepping Stone approach to exploration where human and robotic activities are conducted through progressive expansion outward beyond low- Earth orbit. Results from this study produced a long-term strategy for exploration with near-term implementation plans, program recommendations, and technology investments. Specific results included the development of a common exploration crew vehicle concept, a unified space nuclear strategy, focused bioastronautics research objectives, and an integrated human and robotic exploration strategy. Recommendations from the Exploration Blueprint included the endorsement of the Nuclear Systems Initiative, augmentation of the bioastronautics research, a focused space transportation program including heavy-lift launch and a common exploration vehicle design for ISS and exploration missions, as well as an integrated human and robotic exploration strategy for Mars. Following the results of the Exploration Blueprint study, the NASA Administrator has asked for a recommendation by June, 2003 on the next steps in human and robotic exploration in order to put into context an updated Integrated Space Transportation Plan (post- Columbia) and guide Agency planning. NASA was on the verge of committing significant funding in programs that would be better served if longer term goals were better known including the Orbital Space Plane, research on the ISS, National Aerospace Initiative, Shuttle Life Extension Program, Project Prometheus, as well as a wide range of technology development throughout the Agency. Much of the focus during this period was on integrating the results from the previous studies into more concrete implementation strategies in order to understand the relationship between NASA programs, timing, and resulting budgetary implications. This resulted in an integrated approach including lunar surface operations to retire risk of human Mars missions, maximum use of common and modular systems including what was termed the exploration transfer vehicle, Earth orbit and lunar surface demonstrations of long-life systems, collaboration of human and robotic missions to vastly increase mission return, and high-efficiency transportation systems (nuclear) for deep-space transportation and power. The data provided in this summary viewgraph presentation was developed to begin to address one of the key elements of the emerging implementation strategy, namely how lunar missions help retire risk of human missions to Mars. During this process the scope of the activity broadened into the issue of how testing in general, in various venues including the Moon, can help reduce the risk for Mars missions.

Reducing the Risk of Human Missions to Mars Through Testing

NASA Technical Reports Server (NTRS)

Drake, Bret G.

2007-01-01

During the summer of 2002 the NASA Deputy Administrator charted an internal NASA planning group to develop the rationale for exploration beyond low-Earth orbit. This team, termed the Exploration Blueprint, performed architecture analyses to develop roadmaps for how to accomplish the first steps beyond Low-Earth Orbit through the human exploration of Mars. The previous NASA Exploration Team (NEXT) activities laid the foundation and framework for development of NASA s Integrated Space Plan. The reference missions resulting from the analysis performed by the Exploration Blueprint team formed the basis for requirement definition, systems development, technology roadmapping, and risk assessments for future human exploration beyond low-Earth orbit. Emphasis was placed on developing recommendations on what could be done now to effect future exploration activities. The Exploration Blueprint team embraced the Stepping Stone approach to exploration where human and robotic activities are conducted through progressive expansion outward beyond low- Earth orbit. Results from this study produced a long-term strategy for exploration with near-term implementation plans, program recommendations, and technology investments. Specific results included the development of a common exploration crew vehicle concept, a unified space nuclear strategy, focused bioastronautics research objectives, and an integrated human and robotic exploration strategy. Recommendations from the Exploration Blueprint included the endorsement of the Nuclear Systems Initiative, augmentation of the bioastronautics research, a focused space transportation program including heavy-lift launch and a common exploration vehicle design for ISS and exploration missions, as well as an integrated human and robotic exploration strategy for Mars. Following the results of the Exploration Blueprint study, the NASA Administrator has asked for a recommendation by June, 2003 on the next steps in human and robotic exploration in order to put into context an updated Integrated Space Transportation Plan (post- Columbia) and guide Agency planning. NASA was on the verge of committing significant funding in programs that would be better served if longer term goals were better known including the Orbital Space Plane, research on the ISS, National Aerospace Initiative, Shuttle Life Extension Program, Project Prometheus, as well as a wide range of technology development throughout the Agency. Much of the focus during this period was on integrating the results from the previous studies into more concrete implementation strategies in order to understand the relationship between NASA programs, timing, and resulting budgetary implications. This resulted in an integrated approach including lunar surface operations to retire risk of human Mars missions, maximum use of common and modular systems including what was termed the exploration transfer vehicle, Earth orbit and lunar surface demonstrations of long-life systems, collaboration of human and robotic missions to vastly increase mission return, and high-efficiency transportation systems (nuclear) for deep-space transportation and power. The data provided in this summary viewgraph presentation was developed to begin to address one of the key elements of the emerging implementation strategy, namely how lunar missions help retire risk of human missions to Mars. During this process the scope of the activity broadened into the issue of how testing in general, in various venues including the Moon, can help reduce the risk for Mars missions.

The Efficacy of Radiotherapy in the Treatment of Orbital Pseudotumor

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

Matthiesen, Chance, E-mail: chance-matthiesen@ouhsc.ed; Bogardus, Carl; Thompson, J. Spencer

Purpose: To review institutional outcomes for patients treated with external-beam radiotherapy (EBRT) for orbital pseudotumor. Methods and Materials: This is a single-institution retrospective review of 20 orbits in 16 patients diagnosed with orbital pseudotumor that received EBRT at the University of Oklahoma, Department of Radiation Oncology. Treated patients had a median follow-up of 16.5 months. Results: Fifteen patients (93.7%) were initially treated with corticosteroids. Eight had recurrence after steroid cessation, six were unable to taper corticosteroids completely or partially, and one experienced progression of symptoms despite corticosteroid therapy. Fourteen patients (87.5%) initially responded to radiotherapy indicated by clinical improvement ofmore » preradiation symptoms and/or tapering of corticosteroid dose. Mean EBRT dose was 20 Gy (range, 14-30 Gy). Thirteen patients (81.2%) continued to improve after radiation therapy. Patient outcomes were complete cessation of corticosteroid therapy in nine patients (56.3%) and reduced corticosteroid dose in four patients (25%). Radiotherapy did not achieve long-term control for three patients (18.7%), who still required preradiation corticosteroid dosages. Three patients received retreatment(s) of four orbits, of which two patients achieved long-term symptom control without corticosteroid dependence. One patient received retreatment to an orbit three times, achieving long-term control without corticosteroid dependence. No significant late effects have been observed in retreated patients. Conclusions: Radiotherapy is an effective treatment for acute symptomatic improvement and long-term control of orbital pseudotumor. Orbital retreatment can be of clinical benefit, without apparent increase in morbidity, when initial irradiation fails to achieve complete response.« less

An Entry Flight Controls Analysis for a Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Calhoun, Philip

2000-01-01

The NASA Langley Research Center has been performing studies to address the feasibility of various single-stage to orbit concepts for use by NASA and the commercial launch industry to provide a lower cost access to space. Some work on the conceptual design of a typical lifting body concept vehicle, designated VentureStar(sup TM) has been conducted in cooperation with the Lockheed Martin Skunk Works. This paper will address the results of a preliminary flight controls assessment of this vehicle concept during the atmospheric entry phase of flight. The work includes control analysis from hypersonic flight at the atmospheric entry through supersonic speeds to final approach and landing at subsonic conditions. The requirements of the flight control effectors are determined over the full range of entry vehicle Mach number conditions. The analysis was performed for a typical maximum crossrange entry trajectory utilizing angle of attack to limit entry heating and providing for energy management, and bank angle to modulation of the lift vector to provide downrange and crossrange capability to fly the vehicle to a specified landing site. Sensitivity of the vehicle open and closed loop characteristics to CG location, control surface mixing strategy and wind gusts are included in the results. An alternative control surface mixing strategy utilizing a reverse aileron technique demonstrated a significant reduction in RCS torque and fuel required to perform bank maneuvers during entry. The results of the control analysis revealed challenges for an early vehicle configuration in the areas of hypersonic pitch trim and subsonic longitudinal controllability.

A feasibility study of orbiter flight control experiments

NASA Technical Reports Server (NTRS)

Geissler, W. H.

1978-01-01

The results of a feasibility study of orbiter flight control experiments performed are summarized. Feasibility studies were performed on a group of 14 experiments selected from a candidate list of 35 submitted to the study contractor by the flight control community. Concepts and requirements were developed for the 14 selected experiments and they were ranked on a basis of technical value, feasibility, and cost. It was concluded that all the selected experiments can be considered as potential candidates for the Orbiter Experiment program, which is being formulated for the Orbiter Flight Tests and subsequent operational flights, regardless of the relative ranking established during the study. None of the selected experiments has significant safety implications and the cost of most was estimated to be less than $200K.

Controlled tether extends satellite's orbital range

NASA Astrophysics Data System (ADS)

Wigotsky, V.

1984-06-01

A low orbit satellite tethered to the Space Shuttle Orbiter's cargo bay would be able to conduct upper atmosphere experiments without fear of orbit deterioration. NASA has in light of this initiated a Tethered Satellite System program aimed at the 1987 deployment of a 1,100-lb, 5 ft-diameter satellite to a distance of 6-12 miles from the Space Shuttle on a Kevlar tether. The distance of the fully developed system will be 62 miles, representing an altitude of 80 miles above the earth. Tether diameters under consideration are in the 0.065-0.1 inch range. The satellite control system will consist of a reel drive, a deployment boom, and a boom-mounted tether control, in order to vary tether tension during gravity gradient changes.

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

NASA Astrophysics Data System (ADS)

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.

A recurrent neural-network-based sensor and actuator fault detection and isolation for nonlinear systems with application to the satellite's attitude control subsystem.

PubMed

Talebi, H A; Khorasani, K; Tafazoli, S

2009-01-01

This paper presents a robust fault detection and isolation (FDI) scheme for a general class of nonlinear systems using a neural-network-based observer strategy. Both actuator and sensor faults are considered. The nonlinear system considered is subject to both state and sensor uncertainties and disturbances. Two recurrent neural networks are employed to identify general unknown actuator and sensor faults, respectively. The neural network weights are updated according to a modified backpropagation scheme. Unlike many previous methods developed in the literature, our proposed FDI scheme does not rely on availability of full state measurements. The stability of the overall FDI scheme in presence of unknown sensor and actuator faults as well as plant and sensor noise and uncertainties is shown by using the Lyapunov's direct method. The stability analysis developed requires no restrictive assumptions on the system and/or the FDI algorithm. Magnetorquer-type actuators and magnetometer-type sensors that are commonly employed in the attitude control subsystem (ACS) of low-Earth orbit (LEO) satellites for attitude determination and control are considered in our case studies. The effectiveness and capabilities of our proposed fault diagnosis strategy are demonstrated and validated through extensive simulation studies.

Space shuttle on-orbit flight control software requirements, preliminary version

NASA Technical Reports Server (NTRS)

1975-01-01

Software modules associated with various flight control functions for the space shuttle orbiter are described. Data flow, interface requirements, initialization requirements and module sequencing requirements are considered. Block diagrams and tables are included.

Guidance and Control System for a Satellite Constellation

NASA Technical Reports Server (NTRS)

Bryson, Jonathan Lamar; Cox, James; Mays, Paul Richard; Neidhoefer, James Christian; Ephrain, Richard

2010-01-01

A distributed guidance and control algorithm was developed for a constellation of satellites. The system repositions satellites as required, regulates satellites to desired orbits, and prevents collisions. 1. Optimal methods are used to compute nominal transfers from orbit to orbit. 2. Satellites are regulated to maintain the desired orbits once the transfers are complete. 3. A simulator is used to predict potential collisions or near-misses. 4. Each satellite computes perturbations to its controls so as to increase any unacceptable distances of nearest approach to other objects. a. The avoidance problem is recast in a distributed and locally-linear form to arrive at a tractable solution. b. Plant matrix values are approximated via simulation at each time step. c. The Linear Quadratic Gaussian (LQG) method is used to compute perturbations to the controls that will result in increased miss distances. 5. Once all danger is passed, the satellites return to their original orbits, all the while avoiding each other as above. 6. The delta-Vs are reasonable. The controller begins maneuvers as soon as practical to minimize delta-V. 7. Despite the inclusion of trajectory simulations within the control loop, the algorithm is sufficiently fast for available satellite computer hardware. 8. The required measurement accuracies are within the capabilities of modern inertial measurement devices and modern positioning devices.

ESOC activities during the MIR de-orbit

NASA Astrophysics Data System (ADS)

Klinkrad, H.; Flury, W.; Hernández, C.; Landgraf, M.; Jehn, R.; Christ, U.; Sintoni, F.

2002-11-01

On March 23, 2001, MIR was de-orbited in a controlled fashion, following a successful mission of 15 years. The de-orbiting operations were conducted by the TsUP Mission Control Center, who also consulted entities outside Russia, in order to consolidate their knowledge on the MIR orbit and attitude prior to the initiation of the de-orbit sequence. The European Space Agency ESA through their operations centre ESOC was tasked to support the pre-entry analysis of TsUP by own results, and by routing of Russian and European data via a dedicated communications network. Analysis results produced by ESOC, and details on the data exchange will be highlighted in this paper. The MIR de-orbit and its assessed risk potential will also be compared with the re-entries of Skylab and Salyut-7/Kosmos-1686.

Near Earth Asteroid Rendezvous (NEAR) Revised Eros Orbit Phase Trajectory Design

NASA Technical Reports Server (NTRS)

Helfrich, J; Miller, J. K.; Antreasian, P. G.; Carranza, E.; Williams, B. G.; Dunham, D. W.; Farquhar, R. W.; McAdams, J. V.

1999-01-01

Trajectory design of the orbit phase of the NEAR mission involves a new process that departs significantly from those procedures used in previous missions. In most cases, a precise spacecraft ephemeris is designed well in advance of arrival at the target body. For NEAR, the uncertainty in the dynamic environment around Eros does not allow the luxury of a precise spacecraft trajectory to be defined in advance. The principal cause of this uncertainty is the limited knowledge oi' the gravity field a,-id rotational state of Eros. As a result, the concept for the NEAR trajectory design is to define a number of rules for satisfying spacecraft, mission, and science constraints, and then apply these rules to various assumptions for the model of Eros. Nominal, high, and low Eros mass models are used for testing the trajectory design strategy and to bracket the ranges of parameter variations that are expected upon arrival at the asteroid. The final design is completed after arrival at Eros and determination of the actual gravity field and rotational state. As a result of the unplanned termination of the deep space rendezvous maneuver on December 20, 1998, the NEAR spacecraft passed within 3830 km of Eros on December 23, 1998. This flyby provided a brief glimpse of Eros, and allowed for a more accurate model of the rotational parameters and gravity field uncertainty. Furthermore, after the termination of the deep space rendezvous burn, contact with the spacecraft was lost and the NEAR spacecraft lost attitude control. During the subsequent gyrations of the spacecraft, hydrazine thruster firings were used to regain attitude control. This unplanned thruster activity used Much of the fuel margin allocated for the orbit phase. Consequently, minimizing fuel consumption is now even more important.

Surrogate assisted multidisciplinary design optimization for an all-electric GEO satellite

NASA Astrophysics Data System (ADS)

Shi, Renhe; Liu, Li; Long, Teng; Liu, Jian; Yuan, Bin

2017-09-01

State-of-the-art all-electric geostationary earth orbit (GEO) satellites use electric thrusters to execute all propulsive duties, which significantly differ from the traditional all-chemical ones in orbit-raising, station-keeping, radiation damage protection, and power budget, etc. Design optimization task of an all-electric GEO satellite is therefore a complex multidisciplinary design optimization (MDO) problem involving unique design considerations. However, solving the all-electric GEO satellite MDO problem faces big challenges in disciplinary modeling techniques and efficient optimization strategy. To address these challenges, we presents a surrogate assisted MDO framework consisting of several modules, i.e., MDO problem definition, multidisciplinary modeling, multidisciplinary analysis (MDA), and surrogate assisted optimizer. Based on the proposed framework, the all-electric GEO satellite MDO problem is formulated to minimize the total mass of the satellite system under a number of practical constraints. Then considerable efforts are spent on multidisciplinary modeling involving geosynchronous transfer, GEO station-keeping, power, thermal control, attitude control, and structure disciplines. Since orbit dynamics models and finite element structural model are computationally expensive, an adaptive response surface surrogate based optimizer is incorporated in the proposed framework to solve the satellite MDO problem with moderate computational cost, where a response surface surrogate is gradually refined to represent the computationally expensive MDA process. After optimization, the total mass of the studied GEO satellite is decreased by 185.3 kg (i.e., 7.3% of the total mass). Finally, the optimal design is further discussed to demonstrate the effectiveness of our proposed framework to cope with the all-electric GEO satellite system design optimization problems. This proposed surrogate assisted MDO framework can also provide valuable references for other all-electric spacecraft system design.

CloudSat Anomaly Recovery and Operational Lessons Learned

NASA Technical Reports Server (NTRS)

Witkowski, Mona; Vane, Deborah; Livermore, Thomas; Rokey, Mark; Barthuli, Marda; Gravseth, Ian J.; Pieper, Brian; Rodzinak, Aaron; Silva, Steve; Woznick, Paul;

STS-125 Flight Control Team in WFCR - Orbit 2 - Flight Director Richard LaBrode

NASA Image and Video Library

2009-05-20

JSC2009-E-120845 (20 May 2009) --- The members of the STS-125 Orbit 2 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Rick LaBrode (right) is visible on the front row.

STS-125 Flight Control Team in WFCR - Orbit 3 - Flight Director Paul Dye

NASA Image and Video Library

2009-05-20

JSC2009-E-120846 (20 May 2009) --- The members of the STS-125 Orbit 3 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Paul Dye (center left) is visible on the front row.

STS-131 Flight Control Team in WFCR - Orbit 1 - Flight Director: Richard Jones

NASA Image and Video Library

2010-04-12

JSC2010-E-050680 (12 April 2010) --- The members of the STS-131 Orbit 1 flight control team pose for a group portrait in the space shuttle flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Richard Jones (second left) is on the front row.

The dynamics and control of large flexible space structures, 3. Part A: Shape and orientation control of a platform in orbit using point actuators

NASA Technical Reports Server (NTRS)

Bainum, P. M.; Reddy, A. S. S. R.; Krishna, R.; James, P. K.

1980-01-01

The dynamics, attitude, and shape control of a large thin flexible square platform in orbit are studied. Attitude and shape control are assumed to result from actuators placed perpendicular to the main surface and one edge and their effect on the rigid body and elastic modes is modelled to first order. The equations of motion are linearized about three different nominal orientations: (1) the platform following the local vertical with its major surface perpendicular to the orbital plane; (2) the platform following the local horizontal with its major surface normal to the local vertical; and (3) the platform following the local vertical with its major surface perpendicular to the orbit normal. The stability of the uncontrolled system is investigated analytically. Once controllability is established for a set of actuator locations, control law development is based on decoupling, pole placement, and linear optimal control theory. Frequencies and elastic modal shape functions are obtained using a finite element computer algorithm, two different approximate analytical methods, and the results of the three methods compared.

Modelling and simulation of Space Station Freedom berthing dynamics and control

NASA Technical Reports Server (NTRS)

Cooper, Paul A.; Garrison, James L., Jr.; Montgomery, Raymond C.; Wu, Shih-Chin; Stockwell, Alan E.; Demeo, Martha E.

1994-01-01

A large-angle, flexible, multibody, dynamic modeling capability has been developed to help validate numerical simulations of the dynamic motion and control forces which occur during berthing of Space Station Freedom to the Shuttle Orbiter in the early assembly flights. This paper outlines the dynamics and control of the station, the attached Shuttle Remote Manipulator System, and the orbiter. The simulation tool developed for the analysis is described and the results of two simulations are presented. The first is a simulated maneuver from a gravity-gradient attitude to a torque equilibrium attitude using the station reaction control jets. The second simulation is the berthing of the station to the orbiter with the station control moment gyros actively maintaining an estimated torque equilibrium attitude. The influence of the elastic dynamic behavior of the station and of the Remote Manipulator System on the attitude control of the station/orbiter system during each maneuver was investigated. The flexibility of the station and the arm were found to have only a minor influence on the attitude control of the system during the maneuvers.

Aeroassisted orbital maneuvering using Lyapunov optimal feedback control

NASA Technical Reports Server (NTRS)

Grantham, Walter J.; Lee, Byoung-Soo

1987-01-01

A Liapunov optimal feedback controller incorporating a preferred direction of motion at each state of the system which is opposite to the gradient of a specified descent function is developed for aeroassisted orbital transfer from high-earth orbit to LEO. The performances of the Liapunov controller and a calculus-of-variations open-loop minimum-fuel controller, both of which are based on the 1962 U.S. Standard Atmosphere, are simulated using both the 1962 U.S. Standard Atmosphere and an atmosphere corresponding to the STS-6 Space Shuttle flight. In the STS-6 atmosphere, the calculus-of-variations open-loop controller fails to exit the atmosphere, while the Liapunov controller achieves the optimal minimum-fuel conditions, despite the + or - 40 percent fluctuations in the STS-6 atmosphere.

On the shape and orientation control of an orbiting shallow spherical shell structure

NASA Technical Reports Server (NTRS)

Bainum, P. M.; Reddy, A. S. S. R.

1982-01-01

The dynamics of orbiting shallow flexible spherical shell structures under the influence of control actuators was studied. Control laws are developed to provide both attitude and shape control of the structure. The elastic modal frequencies for the fundamental and lower modes are closely grouped due to the effect of the shell curvature. The shell is gravity stabilized by a spring loaded dumbbell type damper attached at its apex. Control laws are developed based on the pole clustering techniques. Savings in fuel consumption can be realized by using the hybrid shell dumbbell system together with point actuators. It is indicated that instability may result by not including the orbital and first order gravity gradient effects in the plant prior to control law design.

Orbit determination and orbit control for the Earth Observing System (EOS) AM spacecraft

NASA Technical Reports Server (NTRS)

Herberg, Joseph R.; Folta, David C.

1993-01-01

Future NASA Earth Observing System (EOS) Spacecraft will make measurements of the earth's clouds, oceans, atmosphere, land and radiation balance. These EOS Spacecraft will be part of the NASA Mission to Planet Earth. This paper specifically addresses the EOS AM Spacecraft, referred to as 'AM' because it has a sun-synchronous orbit with a 10:30 AM descending node. This paper describes the EOS AM Spacecraft mission orbit requirements, orbit determination, orbit control, and navigation system impact on earth based pointing. The EOS AM Spacecraft will be the first spacecraft to use the TDRSS Onboard Navigation System (TONS) as the primary means of navigation. TONS flight software will process one-way forward Doppler measurements taken during scheduled TDRSS contacts. An extended Kalman filter will estimate spacecraft position, velocity, drag coefficient correction, and ultrastable master oscillator frequency bias and drift. The TONS baseline algorithms, software, and hardware implementation are described in this paper. TONS integration into the EOS AM Spacecraft Guidance, Navigation, and Control (GN&C) System; TONS assisted onboard time maintenance; and the TONS Ground Support System (TGSS) are also addressed.

Application of multi-function display and control technology

NASA Technical Reports Server (NTRS)

Spiger, R. J.; Farrell, R. J.; Holcomb, G. A.

1982-01-01

The NASA orbiter spacecraft incorporates a complex array of systems, displays, and controls. The incorporation of discrete dedicated controls into a multifunction display and control system (MFDCS) offers the potential for savings in weight, power, panel space, and crew training time. Technology identified as applicable to a MFDCS is applied to the orbiter orbital maneuvering system (OMS) and the electrical power distribution and control system (EPDCS) to derive concepts for a MFDCS design. Several concepts of varying degrees of performance and complexity are discussed and a suggested concept for further development is presented in greater detail. Both the hardware and software aspects and the human factors considerations of the designs are included.

A framework for real-time distributed expert systems: On-orbit spacecraft fault diagnosis, monitoring and control

NASA Technical Reports Server (NTRS)

Mullikin, Richard L.

1987-01-01

Control of on-orbit operation of a spacecraft requires retention and application of special purpose, often unique, knowledge of equipment and procedures. Real-time distributed expert systems (RTDES) permit a modular approach to a complex application such as on-orbit spacecraft support. One aspect of a human-machine system that lends itself to the application of RTDES is the function of satellite/mission controllers - the next logical step toward the creation of truly autonomous spacecraft systems. This system application is described.

STS-109 Flight Control Team Photo in WFCR - Orbit 2 with Flight Director Tony Ceccaci.

NASA Image and Video Library

2002-03-05

JSC2002-00574 (5 March 2002) --- The members of the STS-109 Orbit 2 Team pose for a group portrait in the shuttle flight control room (WFCR) in Houston's Mission Control Center (MCC). Flight director Tony Ceccacci holds the STS-109 mission logo.

Optimal Control and Smoothing Techniques for Computing Minimum Fuel Orbital Transfers and Rendezvous

NASA Astrophysics Data System (ADS)

Epenoy, R.; Bertrand, R.

We investigate in this paper the computation of minimum fuel orbital transfers and rendezvous. Each problem is seen as an optimal control problem and is solved by means of shooting methods [1]. This approach corresponds to the use of Pontryagin's Maximum Principle (PMP) [2-4] and leads to the solution of a Two Point Boundary Value Problem (TPBVP). It is well known that this last one is very difficult to solve when the performance index is fuel consumption because in this case the optimal control law has a particular discontinuous structure called "bang-bang". We will show how to modify the performance index by a term depending on a small parameter in order to yield regular controls. Then, a continuation method on this parameter will lead us to the solution of the original problem. Convergence theorems will be given. Finally, numerical examples will illustrate the interest of our method. We will consider two particular problems: The GTO (Geostationary Transfer Orbit) to GEO (Geostationary Equatorial Orbit) transfer and the LEO (Low Earth Orbit) rendezvous.

Hubble Space Telescope Reduced-Gyro Control Law Design, Implementation, and On-Orbit Performance

NASA Technical Reports Server (NTRS)

Clapp, Brian R.; Ramsey, Patrick R.; Wirzburger, John H.; Smith, Daniel C.; VanArsadall, John C.

2008-01-01

Following gyro failures in April 2001 and April 2003, HST Pointing Control System engineers designed reduced-gyro control laws to extend the spacecraft science mission. The Two-Gyro Science (TGS) and One-Gyro Science (OGS) control laws were designed and implemented using magnetometers, star trackers, and Fine Guidance Sensors in succession to control vehicle rate about the missing gyro axes. Both TGS and OGS have demonstrated on-orbit pointing stability of 7 milli-arcseconds or less, which depends upon the guide star magnitude used by the Fine Guidance Sensor. This paper describes the design, implementation, and on-orbit performance of the TGS and OGS control law fine-pointing modes using Fixed Head Star Trackers and Fine Guidance Sensors, after successfully achieving coarse-pointing control using magnetometers.

Overview of Intelligent Power Controller Development for the Deep Space Gateway

NASA Technical Reports Server (NTRS)

Csank, Jeffrey

2017-01-01

Intelligent, or autonomous, control of a spacecraft is an enabling technology that must be developed for deep space human exploration. NASAs current long term human space platform, the International Space Station, which is in Low Earth Orbit, is in almost continuous communication with ground based mission control. This allows near real-time control of all the vehicle core systems, including power, to be controlled by the ground. As focus shifts from Low Earth Orbit, communication time-lag and communication bandwidth limitations beyond geosynchronous orbit does not permit this type of operation. This presentation contains ongoing work at NASA to develop an architecture for autonomous power control and the vehicle manager which monitors, coordinates, and delegates to all the on-board subsystems to enable autonomous control of the complete spacecraft.

Precise orbit determination for the most recent altimeter missions: towards the 1 mm/y stability of the radial orbit error at regional scales

NASA Astrophysics Data System (ADS)

Couhert, Alexandre

The reference Ocean Surface Topography Mission/Jason-2 satellite (CNES/NASA) has been in orbit for six years (since June 2008). It extends the continuous record of highly accurate sea surface height measurements begun in 1992 by the Topex/Poseidon mission and continued in 2001 by the Jason-1 mission. The complementary missions CryoSat-2 (ESA), HY-2A (CNSA) and SARAL/AltiKa (CNES/ISRO), with lower altitudes and higher inclinations, were launched in April 2010, August 2011 and February 2013, respectively. Although the three last satellites fly in different orbits, they contribute to the altimeter constellation while enhancing the global coverage. The CNES Precision Orbit Determination (POD) Group delivers precise and homogeneous orbit solutions for these independent altimeter missions. The focus of this talk will be on the long-term stability of the orbit time series for mean sea level applications on a regional scale. We discuss various issues related to the assessment of radial orbit error trends; in particular orbit errors dependant on the tracking technique, the reference frame accuracy and stability, the modeling of the temporal variations of the geopotential. Strategies are then explored to meet a 1 mm/y radial orbit stability over decadal periods at regional scales, and the challenge of evaluating such an improvement is discussed.

Determination of GPS orbits to submeter accuracy

NASA Technical Reports Server (NTRS)

Bertiger, W. I.; Lichten, S. M.; Katsigris, E. C.

1988-01-01

Orbits for satellites of the Global Positioning System (GPS) were determined with submeter accuracy. Tests used to assess orbital accuracy include orbit comparisons from independent data sets, orbit prediction, ground baseline determination, and formal errors. One satellite tracked 8 hours each day shows rms error below 1 m even when predicted more than 3 days outside of a 1-week data arc. Differential tracking of the GPS satellites in high Earth orbit provides a powerful relative positioning capability, even when a relatively small continental U.S. fiducial tracking network is used with less than one-third of the full GPS constellation. To demonstrate this capability, baselines of up to 2000 km in North America were also determined with the GPS orbits. The 2000 km baselines show rms daily repeatability of 0.3 to 2 parts in 10 to the 8th power and agree with very long base interferometry (VLBI) solutions at the level of 1.5 parts in 10 to the 8th power. This GPS demonstration provides an opportunity to test different techniques for high-accuracy orbit determination for high Earth orbiters. The best GPS orbit strategies included data arcs of at least 1 week, process noise models for tropospheric fluctuations, estimation of GPS solar pressure coefficients, and combine processing of GPS carrier phase and pseudorange data. For data arc of 2 weeks, constrained process noise models for GPS dynamic parameters significantly improved the situation.

Stability and control flight test results of the space transportation system's orbiter

NASA Technical Reports Server (NTRS)

Culp, M. A.; Cooke, D. R.

1982-01-01

Flight testing of the Space Shuttle Orbiter is in progress and current results of the post-flight aerodynamic analyses are discussed. The purpose of these analyses is to reduce the pre-flight aerodynamic uncertainties, thereby leading to operational certification of the Orbiter flight envelope relative to the integrated airframe and flight control system. Primary data reduction is accomplished with a well documented maximum likelihood system identification techniques.

Chaos in navigation satellite orbits caused by the perturbed motion of the Moon

NASA Astrophysics Data System (ADS)

Rosengren, Aaron J.; Alessi, Elisa Maria; Rossi, Alessandro; Valsecchi, Giovanni B.

2015-06-01

Numerical simulations carried out over the past decade suggest that the orbits of the Global Navigation Satellite Systems are unstable, resulting in an apparent chaotic growth of the eccentricity. Here, we show that the irregular and haphazard character of these orbits reflects a similar irregularity in the orbits of many celestial bodies in our Solar system. We find that secular resonances, involving linear combinations of the frequencies of nodal and apsidal precession and the rate of regression of lunar nodes, occur in profusion so that the phase space is threaded by a devious stochastic web. As in all cases in the Solar system, chaos ensues where resonances overlap. These results may be significant for the analysis of disposal strategies for the four constellations in this precarious region of space.

Contingency plans for the ISEE-3 libration-point mission

NASA Technical Reports Server (NTRS)

Dunham, D. W.

1979-01-01

During the planning stage of the International Sun-Earth Explorer-3 (ISEE-3) mission, a recovery strategy was developed in case the Delta rocket underperformed during the launch phase. If a large underburn had occurred, the ISEE-3 spacecraft would have been allowed to complete one revolution of its highly elliptical earth orbit. The recovery plan called for a maneuver near perigee to increase the energy of the off-nominal orbit; a relatively small second maneuver would then insert the spacecraft into a new transfer trajectory toward the desired halo orbit target, and a third maneuver would place the spacecraft in the halo orbit. Results of the study showed that a large range of underburns could be corrected for a total nominal velocity deviation cost within the ISEE-3 fuel budget.

The bistatic radar capabilities of the Medicina radiotelescopes in space debris detection and tracking

NASA Astrophysics Data System (ADS)

Montebugnoli, S.; Pupillo, G.; Salerno, E.; Pluchino, S.; di Martino, M.

2010-03-01

An accurate measurement of the position and trajectory of the space debris fragments is of primary importance for the characterization of the orbital debris environment. The Medicina Radioastronomical Station is a radio observation facility that is here proposed as receiving part of a ground-based space surveillance system for detecting and tracking space debris at different orbital regions (from Low Earth Orbits up to Geostationary Earth Orbits). The proposed system consists of two bistatic radars formed by the existing Medicina receiving antennas coupled with appropriate transmitters. This paper focuses on the current features and future technical development of the receiving part of the observational setup. Outlines of possible transmitting systems will also be given together with the evaluation of the observation strategies achievable with the proposed facilities.

Strengthening the Mars Telecommunications Network

NASA Image and Video Library

2016-11-29

On Nov. 22, 2016, a NASA radio aboard the European Space Agency's (ESA's) Trace Gas Orbiter, which arrived at Mars the previous month, succeeded in its first test of receiving data transmitted from NASA Mars rovers, both Opportunity and Curiosity. This graphic depicts the geometry of Opportunity transmitting data to the orbiter, using the ultra-high frequency (UHF) band of radio wavelengths. The orbiter received that data using one of its twin Electra UHF-band radios. Data that the orbiter's Electra received from the two rovers was subsequently transmitted from the orbiter to Earth, using the orbiter's main X-band radio. The Trace Gas Orbiter is part of ESA's ExoMars program. During the initial months after its Oct. 19, 2016, arrival, it is flying in a highly elliptical orbit. Each loop takes 4.2 days to complete, with distances between the orbiter and the planet's surface ranging from about 60,000 miles (about 100,000 kilometers) to less than 200 miles (less than 310 kilometers). Later, the mission will reshape the orbit to a near-circular path about 250 miles (400 kilometers) above the surface of Mars. Three NASA orbiters and one other ESA orbiter currently at Mars also have relayed data from Mars rovers to Earth. This strategy enables receiving much more data from the surface missions than would be possible with a direct-to-Earth radio link from rovers or stationary landers. Successful demonstration of the capability added by the Trace Gas Orbiter strengthens and extends the telecommunications network at Mars for supporting future missions to the surface of the Red Planet. http://photojournal.jpl.nasa.gov/catalog/PIA21139

Mars Reconnaissance Orbiter Navigation Strategy for Dual Support of Insight and ExoMars Entry, Descent and Landing Demonstrator Module in 2016

NASA Technical Reports Server (NTRS)

Wagner, Sean V.; Menon, Premkumar R.; Chung, Min-Kun J.; Williams, Jessica L.

2015-01-01

Mars Reconnaissance Orbiter (MRO) will support NASA's InSight Mission and ESA's ExoMars Entry, Descent and Landing Demonstrator Module (EDM) in the fall of 2016 when both landers arrive at Mars. MRO provided relay support during the Entry, Descent and Landing (EDL) sequences of Mars Phoenix Lander in 2008 and the Mars Science Laboratory in 2012. Unlike these missions, MRO will coordinate between two EDL events separated by only three weeks: InSight on September 28, 2016 and EDM on October 19, 2016. This paper describes MRO Navigation's maneuver strategy to move MRO's ascending node to meet the In- Sight EDL phasing requirement and support EDM.

Hybrid and Constrained Resolution-of-Identity Techniques for Coulomb Integrals.

PubMed

Duchemin, Ivan; Li, Jing; Blase, Xavier

2017-03-14

The introduction of auxiliary bases to approximate molecular orbital products has paved the way to significant savings in the evaluation of four-center two-electron Coulomb integrals. We present a generalized dual space strategy that sheds a new light on variants over the standard density and Coulomb-fitting schemes, including the possibility of introducing minimization constraints. We improve in particular the charge- or multipole-preserving strategies introduced respectively by Baerends and Van Alsenoy that we compare to a simple scheme where the Coulomb metric is used for lowest angular momentum auxiliary orbitals only. We explore the merits of these approaches on the basis of extensive Hartree-Fock and MP2 calculations over a standard set of medium size molecules.

Advanced control techniques for teleoperation in earth orbit

NASA Technical Reports Server (NTRS)

Bejczy, A. K.; Brooks, T. L.

1980-01-01

Emerging teleoperation tasks in space invite advancements in teleoperator control technology. This paper briefly summarizes the generic issues related to earth orbital applications of teleoperators, and describes teleoperator control technology development work including visual and non-visual sensors and displays, kinesthetic feedback and computer-aided controls. Performance experiments were carried out using sensor and computer aided controls with promising results which are briefly summarized.

Strategies for satellite-based monitoring of CO2 from distributed area and point sources

NASA Astrophysics Data System (ADS)

Schwandner, Florian M.; Miller, Charles E.; Duren, Riley M.; Natraj, Vijay; Eldering, Annmarie; Gunson, Michael R.; Crisp, David

2014-05-01

Atmospheric CO2 budgets are controlled by the strengths, as well as the spatial and temporal variabilities of CO2 sources and sinks. Natural CO2 sources and sinks are dominated by the vast areas of the oceans and the terrestrial biosphere. In contrast, anthropogenic and geogenic CO2 sources are dominated by distributed area and point sources, which may constitute as much as 70% of anthropogenic (e.g., Duren & Miller, 2012), and over 80% of geogenic emissions (Burton et al., 2013). Comprehensive assessments of CO2 budgets necessitate robust and highly accurate satellite remote sensing strategies that address the competing and often conflicting requirements for sampling over disparate space and time scales. Spatial variability: The spatial distribution of anthropogenic sources is dominated by patterns of production, storage, transport and use. In contrast, geogenic variability is almost entirely controlled by endogenic geological processes, except where surface gas permeability is modulated by soil moisture. Satellite remote sensing solutions will thus have to vary greatly in spatial coverage and resolution to address distributed area sources and point sources alike. Temporal variability: While biogenic sources are dominated by diurnal and seasonal patterns, anthropogenic sources fluctuate over a greater variety of time scales from diurnal, weekly and seasonal cycles, driven by both economic and climatic factors. Geogenic sources typically vary in time scales of days to months (geogenic sources sensu stricto are not fossil fuels but volcanoes, hydrothermal and metamorphic sources). Current ground-based monitoring networks for anthropogenic and geogenic sources record data on minute- to weekly temporal scales. Satellite remote sensing solutions would have to capture temporal variability through revisit frequency or point-and-stare strategies. Space-based remote sensing offers the potential of global coverage by a single sensor. However, no single combination of orbit and sensor provides the full range of temporal sampling needed to characterize distributed area and point source emissions. For instance, point source emission patterns will vary with source strength, wind speed and direction. Because wind speed, direction and other environmental factors change rapidly, short term variabilities should be sampled. For detailed target selection and pointing verification, important lessons have already been learned and strategies devised during JAXA's GOSAT mission (Schwandner et al, 2013). The fact that competing spatial and temporal requirements drive satellite remote sensing sampling strategies dictates a systematic, multi-factor consideration of potential solutions. Factors to consider include vista, revisit frequency, integration times, spatial resolution, and spatial coverage. No single satellite-based remote sensing solution can address this problem for all scales. It is therefore of paramount importance for the international community to develop and maintain a constellation of atmospheric CO2 monitoring satellites that complement each other in their temporal and spatial observation capabilities: Polar sun-synchronous orbits (fixed local solar time, no diurnal information) with agile pointing allow global sampling of known distributed area and point sources like megacities, power plants and volcanoes with daily to weekly temporal revisits and moderate to high spatial resolution. Extensive targeting of distributed area and point sources comes at the expense of reduced mapping or spatial coverage, and the important contextual information that comes with large-scale contiguous spatial sampling. Polar sun-synchronous orbits with push-broom swath-mapping but limited pointing agility may allow mapping of individual source plumes and their spatial variability, but will depend on fortuitous environmental conditions during the observing period. These solutions typically have longer times between revisits, limiting their ability to resolve temporal variations. Geostationary and non-sun-synchronous low-Earth-orbits (precessing local solar time, diurnal information possible) with agile pointing have the potential to provide, comprehensive mapping of distributed area sources such as megacities with longer stare times and multiple revisits per day, at the expense of global access and spatial coverage. An ad hoc CO2 remote sensing constellation is emerging. NASA's OCO-2 satellite (launch July 2014) joins JAXA's GOSAT satellite in orbit. These will be followed by GOSAT-2 and NASA's OCO-3 on the International Space Station as early as 2017. Additional polar orbiting satellites (e.g., CarbonSat, under consideration at ESA) and geostationary platforms may also become available. However, the individual assets have been designed with independent science goals and requirements, and limited consideration of coordinated observing strategies. Every effort must be made to maximize the science return from this constellation. We discuss the opportunities to exploit the complementary spatial and temporal coverage provided by these assets as well as the crucial gaps in the capabilities of this constellation. References Burton, M.R., Sawyer, G.M., and Granieri, D. (2013). Deep carbon emissions from volcanoes. Rev. Mineral. Geochem. 75: 323-354. Duren, R.M., Miller, C.E. (2012). Measuring the carbon emissions of megacities. Nature Climate Change 2, 560-562. Schwandner, F.M., Oda, T., Duren, R., Carn, S.A., Maksyutov, S., Crisp, D., Miller, C.E. (2013). Scientific Opportunities from Target-Mode Capabilities of GOSAT-2. NASA Jet Propulsion Laboratory, California Institute of Technology, Pasadena CA, White Paper, 6p., March 2013.

Study of an orbiting tethered dumbbell system having positive orbital energy

NASA Technical Reports Server (NTRS)

Arnold, David A.

1988-01-01

For very long tethered systems the sum of the kinetic and potential energy can be positive. The system remains in a circular orbit as long as the masses remain vertically aligned. The system is unstable without constant control of the alignment. If the upper mass rotates forward in the direction of the orbital motion, the system escapes out of orbit. If the upper mass rotates backward, the system falls out of orbit and the lower mass impacts the body around which the system is orbiting.

Act-and-wait time-delayed feedback control of autonomous systems

NASA Astrophysics Data System (ADS)

Pyragas, Viktoras; Pyragas, Kestutis

2018-02-01

Recently an act-and-wait modification of time-delayed feedback control has been proposed for the stabilization of unstable periodic orbits in nonautonomous dynamical systems (Pyragas and Pyragas, 2016 [30]). The modification implies a periodic switching of the feedback gain and makes the closed-loop system finite-dimensional. Here we extend this modification to autonomous systems. In order to keep constant the phase difference between the controlled orbit and the act-and-wait switching function an additional small-amplitude periodic perturbation is introduced. The algorithm can stabilize periodic orbits with an odd number of real unstable Floquet exponents using a simple single-input single-output constraint control.

Stationkeeping of the First Earth-Moon Libration Orbiters: The ARTEMIS Mission

NASA Technical Reports Server (NTRS)

Folta, David; Woodard, Mark; Cosgrove, D.

2011-01-01

Libration point orbits near collinear locations are inherently unstable and must be controlled. For Acceleration Reconnection and Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) Earth-Moon Lissajous orbit operations, stationkeeping is challenging because of short time scales, large orbital eccentricity of the secondary, and solar gravitational and radiation pressure perturbations. ARTEMIS is the first NASA mission continuously controlled at both Earth-Moon L1 and L2 locations and uses a balance of optimization, spacecraft implementation and constraints, and multi-body dynamics. Stationkeeping results are compared to pre-mission research including mode directions.

Study of orbiter/payload interface communications configuration control board directive from an operational perspective

NASA Technical Reports Server (NTRS)

Addis, A. W.; Tatosian, C. G.; Lidsey, J. F.

1974-01-01

Orbiter/payload data and communications interface was examined. It was found that the Configuration Control Board Directive (CCBD) greatly increases the capability of the orbiter to communicate with a wide variety of projected shuttle payloads. Rather than being derived from individual payload communication requirements, the CCBD appears to be based on an operational philosophy that requires the orbiter to duplicate or augment the ground network/payload communication links. It is suggested that the implementation of the CCBD be reviewed and compared with the Level 1 Program Requirements Document, differences reconciled, and interface characteristics defined.

X33 Reusable Launch Vehicle Control on Sliding Modes: Concepts for a Control System Development

NASA Technical Reports Server (NTRS)

Shtessel, Yuri B.

1998-01-01

Control of the X33 reusable launch vehicle is considered. The launch control problem consists of automatic tracking of the launch trajectory which is assumed to be optimally precalculated. It requires development of a reliable, robust control algorithm that can automatically adjust to some changes in mission specifications (mass of payload, target orbit) and the operating environment (atmospheric perturbations, interconnection perturbations from the other subsystems of the vehicle, thrust deficiencies, failure scenarios). One of the effective control strategies successfully applied in nonlinear systems is the Sliding Mode Control. The main advantage of the Sliding Mode Control is that the system's state response in the sliding surface remains insensitive to certain parameter variations, nonlinearities and disturbances. Employing the time scaling concept, a new two (three)-loop structure of the control system for the X33 launch vehicle was developed. Smoothed sliding mode controllers were designed to robustly enforce the given closed-loop dynamics. Simulations of the 3-DOF model of the X33 launch vehicle with the table-look-up models for Euler angle reference profiles and disturbance torque profiles showed a very accurate, robust tracking performance.

Independent Orbiter Assessment (IOA): Analysis of the Electrical Power Distribution and Control Subsystem, Volume 2

NASA Technical Reports Server (NTRS)

Schmeckpeper, K. R.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Electrical Power Distribution and Control (EPD and C) hardware. The EPD and C hardware performs the functions of distributing, sensing, and controlling 28 volt DC power and of inverting, distributing, sensing, and controlling 117 volt 400 Hz AC power to all Orbiter subsystems from the three fuel cells in the Electrical Power Generation (EPG) subsystem. Volume 2 continues the presentation of IOA analysis worksheets and contains the potential critical items list.

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

NASA Technical Reports Server (NTRS)

Neil, A. L.

1973-01-01

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

Contamination control and cleanliness level integrity for the Space Shuttle Orbiter PLB, payloads and facilities at KSC

NASA Technical Reports Server (NTRS)

Bartelson, D.

1984-01-01

The PLB, its cargo, and payload canister must satisfy the cleanliness requirements of visual clean (VC) level 1, 2, 3, or special as stated in NASA document SN-C-0005A. The specific level of cleanliness is chosen by the payload bay customer for their mission. During orbiter turnaround processing at KSC, the payload bay is exposed to the environments of the Orbiter Processing Facility (OPF) and the Payload Changeout Room (PCR). In supportive response to the orbiter payload bay/facility interface, it is necessary that the facility environment be controlled and monitored to protect the cleanliness/environmental integrity of the payload bay and its cargo. Techniques used to meet environmental requirements during orbiter processing are introduced.

The cislunar low-thrust trajectories via the libration point

NASA Astrophysics Data System (ADS)

Qu, Qingyu; Xu, Ming; Peng, Kun

2017-05-01

The low-thrust propulsion will be one of the most important propulsion in the future due to its large specific impulse. Different from traditional low-thrust trajectories (LTTs) yielded by some optimization algorithms, the gradient-based design methodology is investigated for LTTs in this paper with the help of invariant manifolds of LL1 point and Halo orbit near the LL1 point. Their deformations under solar gravitational perturbation are also presented to design LTTs in the restricted four-body model. The perturbed manifolds of LL1 point and its Halo orbit serve as the free-flight phase to reduce the fuel consumptions as much as possible. An open-loop control law is proposed, which is used to guide the spacecraft escaping from Earth or captured by Moon. By using a two-dimensional search strategy, the ON/OFF time of the low-thrust engine in the Earth-escaping and Moon-captured phases can be obtained. The numerical implementations show that the LTTs achieved in this paper are consistent with the one adopted by the SMART-1 mission.

Robust approximate optimal guidance strategies for aeroassisted orbital transfer missions

NASA Astrophysics Data System (ADS)

Ilgen, Marc R.

This thesis presents the application of game theoretic and regular perturbation methods to the problem of determining robust approximate optimal guidance laws for aeroassisted orbital transfer missions with atmospheric density and navigated state uncertainties. The optimal guidance problem is reformulated as a differential game problem with the guidance law designer and Nature as opposing players. The resulting equations comprise the necessary conditions for the optimal closed loop guidance strategy in the presence of worst case parameter variations. While these equations are nonlinear and cannot be solved analytically, the presence of a small parameter in the equations of motion allows the method of regular perturbations to be used to solve the equations approximately. This thesis is divided into five parts. The first part introduces the class of problems to be considered and presents results of previous research. The second part then presents explicit semianalytical guidance law techniques for the aerodynamically dominated region of flight. These guidance techniques are applied to unconstrained and control constrained aeroassisted plane change missions and Mars aerocapture missions, all subject to significant atmospheric density variations. The third part presents a guidance technique for aeroassisted orbital transfer problems in the gravitationally dominated region of flight. Regular perturbations are used to design an implicit guidance technique similar to the second variation technique but that removes the need for numerically computing an optimal trajectory prior to flight. This methodology is then applied to a set of aeroassisted inclination change missions. In the fourth part, the explicit regular perturbation solution technique is extended to include the class of guidance laws with partial state information. This methodology is then applied to an aeroassisted plane change mission using inertial measurements and subject to uncertainties in the initial value of the flight path angle. A summary of performance results for all these guidance laws is presented in the fifth part of this thesis along with recommendations for further research.

Resistance of CFRP structures to environmental degradation in low Earth orbit

NASA Astrophysics Data System (ADS)

Suliga, Agnieszka

Within this study, a development of a protection strategy for ultra-thin CFRP structures from degrading effects of low Earth orbit (LEO) is presented. The proposed strategy involves an application of a modified epoxy resin system on outer layers of the structure, which is cycloaliphatic in its chemical character and reinforced with POSS nanoparticles. The core of the CFRP structure is manufactured using a highly aromatic epoxy resin system which provides excellent mechanical properties, however, its long-term ageing performance in space is not satisfactory, and hence a surface treatment is required to improve its longevity. The developed resin system presented in this thesis is a hybrid material, designed in such a way that its individual constituents each contribute to combating the detrimental effects of radiation, atomic oxygen (AO), temperature extremes and vacuum induced outgassing of exposed material surfaces while operating in LEO. The cycloaliphatic nature of the outer epoxy increases UV resistance and the embedded silicon nanoparticles improve AO and thermal stability. During the study, a material characterization of the developed cycloaliphatic epoxy resins was performed including the effects of nanoparticles on morphology, curing behaviour, thermal-mechanical properties and surface chemistry. Following on that, the efficacy of the modified resin system on space-like resistance was studied. It was found that when the ultra-thin CFRP structures are covered with the developed resin system, their AO resistance is approximately doubled, UV susceptibility decreased by 80% and thermal stability improved by 20%. Following on the successful launch of the InflateSail mission earlier this year, which demonstrated a sail deployment and a controlled de-orbiting, the findings of this study are of importance for the future generation of similar, but significantly longer missions. Ensuring resistance of CFRP structures in a highly corrosive LEO environment is a critical requirement to make their use in space applications truly feasible.

Highly efficient perturbative + variational strategy based on orthogonal valence bond theory for the evaluation of magnetic coupling constants. Application to the trinuclear Cu(ii) site of multicopper oxidases.

PubMed

Tenti, Lorenzo; Maynau, Daniel; Angeli, Celestino; Calzado, Carmen J

2016-07-21

A new strategy based on orthogonal valence-bond analysis of the wave function combined with intermediate Hamiltonian theory has been applied to the evaluation of the magnetic coupling constants in two AF systems. This approach provides both a quantitative estimate of the J value and a detailed analysis of the main physical mechanisms controlling the coupling, using a combined perturbative + variational scheme. The procedure requires a selection of the dominant excitations to be treated variationally. Two methods have been employed: a brute-force selection, using a logic similar to that of the CIPSI approach, or entanglement measures, which identify the most interacting orbitals in the system. Once a reduced set of excitations (about 300 determinants) is established, the interaction matrix is dressed at the second-order of perturbation by the remaining excitations of the CI space. The diagonalization of the dressed matrix provides J values in good agreement with experimental ones, at a very low-cost. This approach demonstrates the key role of d → d* excitations in the quantitative description of the magnetic coupling, as well as the importance of using an extended active space, including the bridging ligand orbitals, for the binuclear model of the intermediates of multicopper oxidases. The method is a promising tool for dealing with complex systems containing several active centers, as an alternative to both pure variational and DFT approaches.

Investigation of electrodynamic stabilization and control of long orbiting tethers

NASA Technical Reports Server (NTRS)

Colombo, G.; Arnold, D.

1984-01-01

The state-of-the-art in tether modelling among participants in the Tethered Satellite System (TSS) Program, the slack tether and its behavior, and certain advanced applications of the tether to problems in orbital mechanics are identified. The features and applications of the TSS software set are reviewed. Modelling the slack tether analytically with as many as 50 mass points and the application of this new model to a study of the behavior of a broken tether near the Shuttle are described. A reel control algorithm developed by SAO and examples of its use are described, including an example which also demonstrates the use of the tether in transferring a heavy payload from a low-orbiting Shuttle to a high circular orbit. Capture of a low-orbiting payload by a Space Station in high circular orbit is described. Energy transfer within a dumbbell-type spacecraft by cyclical reeling operations or gravitational effects on the natural elasticity of the connecting tether, it is shown, can circularize the orbit of the spacecraft.

An Optimized Method to Detect BDS Satellites' Orbit Maneuvering and Anomalies in Real-Time.

PubMed

Huang, Guanwen; Qin, Zhiwei; Zhang, Qin; Wang, Le; Yan, Xingyuan; Wang, Xiaolei

2018-02-28

The orbital maneuvers of Global Navigation Satellite System (GNSS) Constellations will decrease the performance and accuracy of positioning, navigation, and timing (PNT). Because satellites in the Chinese BeiDou Navigation Satellite System (BDS) are in Geostationary Orbit (GEO) and Inclined Geosynchronous Orbit (IGSO), maneuvers occur more frequently. Also, the precise start moment of the BDS satellites' orbit maneuvering cannot be obtained by common users. This paper presented an improved real-time detecting method for BDS satellites' orbit maneuvering and anomalies with higher timeliness and higher accuracy. The main contributions to this improvement are as follows: (1) instead of the previous two-steps method, a new one-step method with higher accuracy is proposed to determine the start moment and the pseudo random noise code (PRN) of the satellite orbit maneuvering in that time; (2) BDS Medium Earth Orbit (MEO) orbital maneuvers are firstly detected according to the proposed selection strategy for the stations; and (3) the classified non-maneuvering anomalies are detected by a new median robust method using the weak anomaly detection factor and the strong anomaly detection factor. The data from the Multi-GNSS Experiment (MGEX) in 2017 was used for experimental analysis. The experimental results and analysis showed that the start moment of orbital maneuvers and the period of non-maneuver anomalies can be determined more accurately in real-time. When orbital maneuvers and anomalies occur, the proposed method improved the data utilization for 91 and 95 min in 2017.

An Optimized Method to Detect BDS Satellites’ Orbit Maneuvering and Anomalies in Real-Time

PubMed Central

Huang, Guanwen; Qin, Zhiwei; Zhang, Qin; Wang, Le; Yan, Xingyuan; Wang, Xiaolei

2018-01-01

The orbital maneuvers of Global Navigation Satellite System (GNSS) Constellations will decrease the performance and accuracy of positioning, navigation, and timing (PNT). Because satellites in the Chinese BeiDou Navigation Satellite System (BDS) are in Geostationary Orbit (GEO) and Inclined Geosynchronous Orbit (IGSO), maneuvers occur more frequently. Also, the precise start moment of the BDS satellites’ orbit maneuvering cannot be obtained by common users. This paper presented an improved real-time detecting method for BDS satellites’ orbit maneuvering and anomalies with higher timeliness and higher accuracy. The main contributions to this improvement are as follows: (1) instead of the previous two-steps method, a new one-step method with higher accuracy is proposed to determine the start moment and the pseudo random noise code (PRN) of the satellite orbit maneuvering in that time; (2) BDS Medium Earth Orbit (MEO) orbital maneuvers are firstly detected according to the proposed selection strategy for the stations; and (3) the classified non-maneuvering anomalies are detected by a new median robust method using the weak anomaly detection factor and the strong anomaly detection factor. The data from the Multi-GNSS Experiment (MGEX) in 2017 was used for experimental analysis. The experimental results and analysis showed that the start moment of orbital maneuvers and the period of non-maneuver anomalies can be determined more accurately in real-time. When orbital maneuvers and anomalies occur, the proposed method improved the data utilization for 91 and 95 min in 2017. PMID:29495638

Boeing Low-Thrust Geosynchronous Transfer Mission Experience

NASA Technical Reports Server (NTRS)

Poole, Mark; Ho, Monte

2007-01-01

Since 2000, Boeing 702 satellites have used electric propulsion for transfer to geostationary orbits. The use of the 25cm Xenon Ion Propulsion System (25cm XIPS) results in more than a tenfold increase in specific impulse with the corresponding decrease in propellant mass needed to complete the mission when compared to chemical propulsion[1]. In addition to more favorable mass properties, with the use of XIPS, the 702 has been able to achieve orbit insertions with higher accuracy than it would have been possible with the use of chemical thrusters. This paper describes the experience attained by using the 702 XIPS ascent strategy to transfer satellite to geosynchronous orbits.

Space Station automated systems testing/verification and the Galileo Orbiter fault protection design/verification

NASA Technical Reports Server (NTRS)

Landano, M. R.; Easter, R. W.

1984-01-01

Aspects of Space Station automated systems testing and verification are discussed, taking into account several program requirements. It is found that these requirements lead to a number of issues of uncertainties which require study and resolution during the Space Station definition phase. Most, if not all, of the considered uncertainties have implications for the overall testing and verification strategy adopted by the Space Station Program. A description is given of the Galileo Orbiter fault protection design/verification approach. Attention is given to a mission description, an Orbiter description, the design approach and process, the fault protection design verification approach/process, and problems of 'stress' testing.

Mars Geoscience Climatology Orbiter (MGCO) extended study: Technical volume

NASA Technical Reports Server (NTRS)

1983-01-01

The FLTSATCOM Earth orbiting communications satellite is a prominent candidate to serve as the Mars Geoscience Climatology Orbiter (MGCO) spacecraft. Major aspects directly applicable are: (1) the incorporation of solid orbit insertion motor; (2) the ability to cruise to Mars in the spin-stabilized mode; (3) ample capability for payload mass and power; (4) attitude control tried to nadir and orbit plane coordinates; (5) exemplary Earth orbital performance record and projected lifetime; and (6) existence of an on-going procurement into the MGCO time period.

ISS15A Flight Control Team in FCR-1 Orbit 1 - Flight Director Kwatsi Alibaruho

NASA Image and Video Library

2009-03-20

JSC2009-E-060959 (20 March 2009) --- The members of the STS-119/15A ISS Orbit 1 flight control team pose for a group portrait in the space station flight control room in the Mission Control Center at NASA’s Johnson Space Center. Flight director Kwatsi Alibaruho (right) is visible on the front row.

STS-131/19A Flight Control Team in FCR-1 - Orbit 1- Flight Director Ron Spencer

NASA Image and Video Library

2010-04-14

JSC2010-E-052008 (14 April 2010) --- The members of the STS-131/19A ISS Orbit 2 flight control team pose for a group portrait in the space station flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Ron Spencer (right) holds the STS-131 mission logo.

ISS ULF2 Flight Control Team in FCR-1 - Orbit 3 - Flight Director David Korth

NASA Image and Video Library

2009-03-20

JSC2009-E-061164 (20 March 2009) --- The members of the STS-119/15A ISS Orbit 3 flight control team pose for a group portrait in the space station flight control room in the Mission Control Center at NASA’s Johnson Space Center. Flight director David Korth (right) is visible on the front row.

STS-131/19A Flight Control Team in FCR-1 - Orbit 1- Flight Director Courtney McMillan

NASA Image and Video Library

2010-04-14

JSC2010-E-052979 (14 April 2010) --- The members of the STS-131/19A ISS Orbit 1 flight control team pose for a group portrait in the space station flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Courtenay McMillan (center) stands on the front row.

STS-131/19A Flight Control Team in FCR-1 - Orbit 3- Flight Director Ed Van Cise

NASA Image and Video Library

2010-04-14

JSC2010-E-052556 (14 April 2010) --- The members of the STS-131/19A ISS Orbit 3 flight control team pose for a group portrait in the space station flight control room in the Mission Control Center at NASA's Johnson Space Center. Flight director Ed Van Cise holds the STS-131 mission logo.

Quantifying and Improving International Space Station Survivability Following Orbital Debris Penetration

NASA Technical Reports Server (NTRS)

Williamsen, Joel; Evans, Hilary; Bohl, Bill; Evans, Steven; Parker, Nelson (Technical Monitor)

2001-01-01

The increase of the orbital debris environment in low-earth orbit has prompted NASA to develop analytical tools for quantifying and lowering the likelihood of crew loss following orbital debris penetration of the International Space Station (ISS). NASA uses the Manned Spacecraft and Crew Survivability (MSCSurv) computer program to simulate the events that may cause crew loss following orbital debris penetration of ISS manned modules, including: (1) critical cracking (explosive decompression) of the module; (2) critical external equipment penetration (such as hydrazine and high pressure tanks); (3) critical internal system penetration (guidance, control, and other vital components); (4) hazardous payload penetration (furnaces, pressure bottles, and toxic substances); (5) crew injury (from fragments, overpressure, light flash, and temperature rise); (6) hypoxia from loss of cabin pressure; and (7) thrust from module hole causing high angular velocity (occurring only when key Guidance, Navigation, and Control (GN&C) equipment is damaged) and, thus, preventing safe escape vehicle (EV) departure. MSCSurv is also capable of quantifying the 'end effects' of orbital debris penetration, such as the likelihood of crew escape, the probability of each module depressurizing, and late loss of station control. By quantifying these effects (and their associated uncertainties), NASA is able to improve the likelihood of crew survivability following orbital debris penetration due to improved crew operations and internal designs.

Space Operations Learning Center (SOLC) iPhone/iPad Application

NASA Technical Reports Server (NTRS)

Binebrink, Daniel; Kuok, Heng; Hammond, Malinda; Hull, Scott

2013-01-01

This iPhone application, Space Junk Sammy, is intended to be an educational application designed for Apple iPhones and iPads. This new concept educates kids in an innovative way about how orbital debris affects space missions. Orbital debris is becoming a very significant concern for NASA and all Earthorbiting space missions. Spacecraft in low-Earth orbit are in constant danger of being potentially damaged or destroyed by debris. High-profile spacecraft such as the International Space Station (ISS) and Hubble Space Telescope are dealing with orbital debris on a regular basis. Other basic educational concepts that are portrayed are low-Earth orbits, satellites, ISS, attitude control, and other facts that can be presented in betweenlevel popup screens. The Orbital Debris Cleanup game is relatively simple from the user s technical standpoint. It is a 2D game where the user s avatar is a satellite buddy, named Sammy, in orbit around Earth. Sammy is controlled by the user with the device s gyroscope as well as touchscreen controls. It has equipment used for taking care of the space debris objects on the screen. Sammy also has a claw, a laser deflector, and hydrazine rockets to grab or push the debris objects into a higher orbit or into a lower orbit to burn up in the Earth s atmosphere. The user interface shows Sammy and space debris objects constantly moving from left to right, where Sammy is trying to catch the debris objects before they move off the right side of the screen. Everything will be in constant motion to increase fun and add to the realism of orbiting the Earth. The satellite buddy is used to clean up the space debris and protect other satellites. Later levels will include a laser deflector and hydrazine rockets instead of a robotic claw to push the orbital debris into a higher orbit and out of the path of other satellites

Proving the Space Transportation System: the Orbital Flight Test Program

NASA Technical Reports Server (NTRS)

Reichhardt, T.

1982-01-01

The main propulsion system, solid rocket boosters, external tank, orbital maneuvering system, spacecraft orbital operations (thermal tests, attitude control and remote manipulator), and return to Earth are outlined for the first four STS missions.

Automated Orbit Determination System (AODS) requirements definition and analysis

NASA Technical Reports Server (NTRS)

Waligora, S. R.; Goorevich, C. E.; Teles, J.; Pajerski, R. S.

1980-01-01

The requirements definition for the prototype version of the automated orbit determination system (AODS) is presented including the AODS requirements at all levels, the functional model as determined through the structured analysis performed during requirements definition, and the results of the requirements analysis. Also specified are the implementation strategy for AODS and the AODS-required external support software system (ADEPT), input and output message formats, and procedures for modifying the requirements.