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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 163 records · Page 9

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NASA NTRS1996

Exact-Output Tracking Theory for Systems with Parameter Jumps

In this paper we consider the exact output tracking problem for systems with parameter jumps. Necessary and sufficient conditions are derived for the elimination of switching-introduced output transient. Previous works have studied this problem by developing a regulator that maintains exact tracking through parameter jumps (switches). Such techniques are, however, only applicable to minimum-phase systems. In contrast, our approach is applicable to nonminimum-phase systems and obtains bounded but possibly non-causal solutions. If the reference trajectories are generated by an exo-system, then we develop an exact-tracking controller in a feedback form. As in standard regulator theory, we obtain a linear map from the states of the exo-system to the desired system state which is defined via a matrix differential equation. The constant solution of this differential equation provides asymptotic tracking, and coincides with the feedback law used in standard regulator theory. The obtained results are applied to a simple flexible manipulator with jumps in the pay-load mass.

Devasia, Santosh↗

NASA NTRS1997

Exact-Output Tracking Theory for Systems with Parameter Jumps

We consider the exact output tracking problem for systems with parameter jumps. Necessary and sufficient conditions are derived for the elimination of switching-introduced output transient. Previous works have studied this problem by developing a regulator that maintains exact tracking through parameter jumps (switches). Such techniques are, however, only applicable to minimum-phase systems. In contrast, our approach is applicable to non-minimum-phase systems and it obtains bounded but possibly non-causal solutions. If the reference trajectories are generated by an exosystem, then we develop an exact-tracking controller in a feed-back form. As in standard regulator theory, we obtain a linear map from the states of the exosystem to the desired system state which is defined via a matrix differential equation. The constant solution of this differential equation provides asymptotic tracking, and coincides with the feedback law used in standard regulator theory. The obtained results are applied to a simple flexible manipulator with jumps in the pay-load mass.

Devasia, Santosh↗

NASA NTRS1999

Feasibility of Decentralized Linear-Quadratic-Gaussian Control of Autonomous Distributed Spacecraft

A distributed satellite formation, modeled as an arbitrary number of fully connected nodes in a network, could be controlled using a decentralized controller framework that distributes operations in parallel over the network. For such problems, a solution that minimizes data transmission requirements, in the context of linear-quadratic-Gaussian (LQG) control theory, was given by Speyer. This approach is advantageous because it is non-hierarchical, detected failures gracefully degrade system performance, fewer local computations are required than for a centralized controller, and it is optimal with respect to the standard LQG cost function. Disadvantages of the approach are the need for a fully connected communications network, the total operations performed over all the nodes are greater than for a centralized controller, and the approach is formulated for linear time-invariant systems. To investigate the feasibility of the decentralized approach to satellite formation flying, a simple centralized LQG design for a spacecraft orbit control problem is adapted to the decentralized framework. The simple design uses a fixed reference trajectory (an equatorial, Keplerian, circular orbit), and by appropriate choice of coordinates and measurements is formulated as a linear time-invariant system.

Carpenter, J. Russell↗

NASA NTRS1999

Feasibility of Decentralized Linear-Quadratic-Gaussian Control of Autonomous Distributed Spacecraft

Carpenter, J. Russell↗

NASA NTRS2002

Improvements in Thermal Protection Sizing Capabilities for TCAT: Conceptual Design for Advanced Space Transportation Systems

The Thermal Calculation Analysis Tool (TCAT), originally developed for the Space Systems Design Lab at the Georgia Institute of Technology, is a conceptual design tool capable of integrating aeroheating analysis into conceptual reusable launch vehicle design. It provides Thermal Protection System (TPS) unit thicknesses and acreage percentages based on the geometry of the vehicle and a reference trajectory to be used in calculation of the total cost and weight of the vehicle design. TCAT has proven to be reasonably accurate at calculating the TPS unit weights for in-flight trajectories; however, it does not have the capability of sizing TPS materials above cryogenic fuel tanks for ground hold operations. During ground hold operations, the vehicle is held for a brief period (generally about two hours) during which heat transfer from the TPS materials to the cryogenic fuel occurs. If too much heat is extracted from the TPS material, the surface temperature may fall below the freezing point of water, thereby freezing any condensation that may be present at the surface of the TPS. Condensation or ice on the surface of the vehicle is potentially hazardous to the mission and can also damage the TPS. It is questionable whether or not the TPS thicknesses provided by the aeroheating analysis would be sufficiently thick to insulate the surface of the TPS from the heat transfer to the fuel. Therefore, a design tool has been developed that is capable of sizing TPS materials at these cryogenic fuel tank locations to augment TCAT's TPS sizing capabilities.

Olds, John R.↗

NASA NTRS1961

Study of the Use of a Terminal Controller Technique for Reentry Guidance of a Capsule-Type Vehicle

A study has been made of the use o f a terminal controller technique i n the guidance of a high-drag, variable-lift reentry vehicle to a desired landing point. The technique uses linearized equations of motion attained by the perturbation of the dependent variables from those of a reference trajectory. The guidance system continuously predicts the terminal range error and uses this error to control the angle of attack of the vehicle in an on-off manner until the predicted range error is within +-O.1 degrees of the required arc or +-6.9 miles.

Foudriat, Edwin C.↗

NASA NTRS2006

MSL EDL Entry Guidance using the Entry Terminal Point Controller

The Mars Science Laboratory will be the first Mars mission to attempt a guided entry with the objective of safely delivering the entry vehicle to a survivable parachute deploy state within 10 km of the pre-designated landing site. The Entry Terminal Point Controller guidance algorithm is derived from the final phase Apollo Command Module guidance and, like Apollo, modulates the bank angle to control range based on deviations in range, altitude rate, and drag acceleration from a reference trajectory. For application to Mars landers which must make use of the tenuous Martian atmosphere, it is critical to balance the lift of the vehicle to minimize the range while still ensuring a safe deploy altitude. An overview of the process to generate optimized guidance settings is presented, discussing improvements made over the last four years. Performance tradeoffs between ellipse size and deploy altitude will be presented, along with imposed constraints of entry acceleration and heating. Performance sensitivities to the bank reversal deadbands, heading alignment, attitude initialization error, and atmospheric delivery errors are presented. Guidance settings for contingency operations, such as those appropriate for severe dust storm scenarios, are evaluated.

Source record↗

NASA NTRS2002

A survey of spacecraft formation flying guidance

This paper is the first comprehensive spacecraft formation flying guidance (FFG) survey. Here by the term guidance we mean both path planning (i. e., reference trajectory generation) and optimal, open loop control design. FFG naturally divides into two areas: Deep Space (DS), in which relative spacecraft dynamics reduce to double integrator form, and Planetary Orbital Environments (POE), in which they do not (e.9. libration point formations). Both areas consider optimal formation reconfigurations. In addition, DS FFG addresses optimal u, v-coverages for multiple spacecraft interferometers and rest-to-rest rotations. The main focus of the POE literature, however, is 'assive apertures.' These are periodic and fuel-eficient relative spacecraft trajectories that accomplish scientific objectives (e.9. synthesizing an aperture).

formation flying guidance optimal control path pla↗

NASA NTRS2003

Cassini tour navigation strategy

The Cassini-Huygens spacecraft was launched on October 15, 1997 as a joint NASA/ESA mission to explore Saturn. After a 7 year cruise the spacecraft will enter orbit around Saturn on 1 July 2004 for a 4 year investigation of the Saturnian system. The Cassini Navigation Team is responsible for designing the reference trajectory and conducting operations to realize this design. This paper describes the strategy for achieving project requirements, the characteristics of the Cassini navigation challenge, and the underlying assumptions.

Saturn↗

NASA NTRS1989

Method and apparatus for adaptive force and position control of manipulators

The present invention discloses systematic methods and apparatus for the design of real time controllers. Real-time control employs adaptive force/position by use of feedforward and feedback controllers, with the feedforward controller being the inverse of the linearized model of robot dynamics and containing only proportional-double-derivative terms is disclosed. The feedback controller, of the proportional-integral-derivative type, ensures that manipulator joints follow reference trajectories and the feedback controller achieves robust tracking of step-plus-exponential trajectories, all in real time. The adaptive controller includes adaptive force and position control within a hybrid control architecture. The adaptive controller, for force control, achieves tracking of desired force setpoints, and the adaptive position controller accomplishes tracking of desired position trajectories. Circuits in the adaptive feedback and feedforward controllers are varied by adaptation laws.

Seraji, Homayoun↗

NASA NTRS2007

Optimisation of the Future Routine Orbit for Mars Express

Mars Express (MEX), the first planetary mission of the European Space Agency (ESA), reached Mars on December 25th 2003. Since then it is performing routine operations. Its operational phase had to cover one Martian year, with the possibility of an extension for a second Martian year (i.e. until November 2007). The end of the mission extension is approaching but, given the good health of the payload instruments and the high science return of the mission, there is a strong will to achieve further extensions. Mars Express is also seen as an important asset, capable to provide relay functions for future Martian missions. The ESA Science Program Committee has recently approved a second extension of the MEX mission until May 2009 and even further extensions are possible. Mars Express has an eccentric quasi-polar orbit with a period of approximately 6.72 hours and a pericentre height of about 300 km. Science observations are mainly performed at pericentre (but not only). In addition the orbit has a resonance of 11 revolutions per 3 Martian days. This means that ground tracks corresponding to orbits separated by 11 revolutions are adjacent, such that a given area can be covered by the on-board camera without leaving gaps. The J2 effect of Mars causes a drift of both ascending node and argument of pericentre. The drift of argument of pericentre makes it possible to observe periodically all Mars latitudes from close distance. Illumination conditions at pericentre are influenced by both the drift of the argument of pericentre and the drift of ascending node, as well as by the rotation of Mars around the Sun. The original MEX routine orbit was optimized for the duration of the nominal mission and extension, such that it produced a balanced share of day-side observations (for the optical instruments) and night-side observations (for the radar). The orbit was thus not optimized for the time beyond the assumed extension. Indeed, the evolution of the ascending node and argument of pericentre would cause in the following years a drift of the pericentre towards night-side observation conditions, hence uninteresting for the optical instruments. In order to prevent this an optimisation process for the future routine orbit has taken place. The share between day-side and night-side observations can be controlled by adjusting the drift of argument of pericentre and ascending node. This can in particular be achieved by changing the semimajor axis, eccentricity and/or inclination. A change of inclination is inefficient compared to a change in semimajor axis and eccentricity, and has therefore been discarded. An in-plane maneuvre can be performed to change both semi-major axis and eccentricity, and thus the period of the orbit. Although an apocentre manoeuvre is cheaper in terms of deltaV, it would result in raising the pericentre height, which is unfavourable for close observations. Hence a pericentre manoeuvre is proposed, which will increase the apocentre height. A repeat cycle is still required to allow mapping areas with adjacent ground tracks, so the change of semimajor axis must result in a new resonance. Resonances 18:5, 25:7 and 7:2 have been considered as potential candidates. The resulting long term evolution of the observation conditions has been analysed. Finally it has been decided to perform a change of orbit to reach the 18:5. Another aspect of the optimisation process is the control of the ground track. The previous MEX reference trajectory included regular maneuvres at every apocentre in order to adjust the orbital period, such that the separation of the ground tracks would be optimal, regardless of the latitude of pericentre. The implementation of the actual delatVs on-board was done partly by optimizing the attitude of reaction-wheel desaturation activities. Despite of it, this strategy has a significant propellant cost, because it prevents to optimize reaction wheel de-saturation activities to minimize propellant consumption. Therore, with the aim at preserving propellant resources for a long time extension it has been agreed to stop the ground-track control. This requires now a more accurate science operation planning, with improved attitude pointing control. Finally, the approach to phase Mars Express to provide back-up relay functions for NASA Phoenix landing is explained. In the context of the routine trajectory optimisation a new requirement for close fly-bys at Phobos, with different observation geometries, has been specified. The approach to fulfill this requirement is explained.

Carranza, Manuel↗

NASA NTRS2010

XMbodyinfo

XMbodyinfo was designed to evaluate potential reference trajectories, providing a proficient way to assess the quality of all satellite body flybys for a Cassini type mission tour. It is autonomous and will generate a variety of ORS (optical remote sensing) and FPW (fields, particles, and waves) plots that aid in the evaluation, qualification, selection, and improvement of a potential tour.

Roumeliotis, Chris↗

NASA NTRS2011

Mars Science Laboratory Entry Guidance

The Mars Science Laboratory will be the first Mars mission to attempt a guided entry with the objective of safely delivering the entry vehicle to a survivable parachute deploy state within 12.5 km of the pre-designated parachute deploy coordinates. The Entry Terminal Point Controller guidance algorithm is derived from the final phase Apollo Command Module guidance and, like Apollo, modulates the bank angle to control range based on deviations in range, altitude rate, and drag acceleration from a reference trajectory. For application to Mars landers which must make use of the tenuous Martian atmosphere, it is critical to balance the lift of the vehicle to minimize the range while still ensuring a safe deploy altitude. An overview of the process to generate optimized guidance settings is presented, discussing improvements made over the last nine years. Performance tradeoffs between ellipse size and deploy altitude will be presented, along with imposed constraints of entry acceleration and heating. Performance sensitivities to the bank reversal deadbands, heading alignment, attitude initialization error, and entry delivery errors are presented.

Mendeck, Gavin F.↗

NASA NTRS2011

Mars Science Laboratory Entry Guidance Improvements for Mars 2018 (DRAFT)

In 2011, the Mars Science Laboratory (MSL) will be launched in a mission to deliver the largest and most capable rover to date to the surface of Mars. A follow on MSL-derived mission, referred to as Mars 2018, is planned for 2018. Mars 2018 goals include performance enhancements of the Entry, Descent and Landing over that of its predecessor MSL mission of 2011. This paper will discuss the main elements of the modified 2018 EDL preliminary design that will increase performance on the entry phase of the mission. In particular, these elements will increase the parachute deploy altitude to allow for more time margin during the subsequent descent and landing phases and reduce the delivery ellipse size at parachute deploy through modifications in the entry reference trajectory design, guidance trigger logic design, and the effect of additional navigation hardware.

Garcia-Llama, Eduardo↗

NASA NTRS2012

Post-Flight EDL Entry Guidance Performance of the 2011 Mars Science Laboratory Mission

The 2011 Mars Science Laboratory was the first successful Mars mission to attempt a guided entry which safely delivered the rover to a final position approximately 2 km from its target within a touchdown ellipse of 19.1 km x 6.9 km. The Entry Terminal Point Controller guidance algorithm is derived from the final phase Apollo Command Module guidance and, like Apollo, modulates the bank angle to control the range flown. For application to Mars landers which must make use of the tenuous Martian atmosphere, it is critical to balance the lift of the vehicle to minimize the range error while still ensuring a safe deploy altitude. An overview of the process to generate optimized guidance settings is presented, discussing improvements made over the last nine years. Key dispersions driving deploy ellipse and altitude performance are identified. Performance sensitivities including attitude initialization error and the velocity of transition from range control to heading alignment are presented. Just prior to the entry and landing of MSL in August 2012, the EDL team examined minute tuning of the reference trajectory for the selected landing site, analyzed whether adjustment of bank reversal deadbands were necessary, the heading alignment velocity trigger was in union with other parameters to balance the EDL risks, and the vertical L/D command limits. This paper details a preliminary postflight assessment of the telemetry and trajectory reconstruction that is being performed, and updates the information presented in the former paper Entry Guidance for the 2011 Mars Science Laboratory Mission (AIAA Atmospheric Flight Mechanics Conference; 8-11 Aug. 2011; Portland, OR; United States)

Mendeck, Gavin F.↗

NASA NTRS2013

Explicit Low-Thrust Guidance for Reference Orbit Targeting

The problem of a low-thrust spacecraft controlled to a reference orbit is addressed in this paper. A simple and explicit low-thrust guidance scheme with constrained thrust magnitude is developed by combining the fundamental equations of motion for constrained systems from analytical dynamics with a Lyapunov-based method. Examples are given for a spacecraft controlled to a reference trajectory in the circular restricted three body problem.

low-thrust guidance↗

NASA NTRS2014

Orbit Determination Covariance Analysis for the Europa Clipper Mission

A new Jovian satellite tour is proposed by NASA, which would include numerous flybys of the moon Europa, and would explore its potential habitability by characterizing the existence of any water within and beneath Europa's ice shell. This paper describes the results of a covariance study that was undertaken on a sample tour to assess the navigational challenges and capabilities of such a mission from an orbit determination (OD) point of view, and to help establish a delta V budget for the maneuvers needed to keep the spacecraft on the reference trajectory. Additional parametric variations from the baseline case were also investigated. The success of the Europa Clipper mission will depend on the science measurements that it will enable. Meeting the requirements of the instruments onboard the spacecraft is an integral part of this analysis.

satellite tour↗

NASA NTRS2016

Periodic Orbit-Attitude Solutions in the Planar Elliptic Restricted Three-Body Problem

The pitch motion of a spacecraft in the planar elliptic restricted three-body system is studied. Previous studies laid the foundation for spacecraft stability analysis with a small perturbation to the zero pitch motion. In this study, a cell mapping approach that combines analytical and numerical techniques is used to study the global behavior of the full nonlinear spacecraft attitude in which coupling between orbital dynamics and attitude occurs. The effect of gravity gradient torques, orbital eccentricity, and the spacecraft configuration at different Lagrangian points is analyzed. Multiple-period periodic solutions and invariant surfaces are presented for different cases. Reference trajectories around the Lagrangian points are also considered to study coupled dynamics.

Koh, Dayung↗