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Onboard Burn targeting and Guidance of NASA Human Flight Vehicles

JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2024年第11期47卷 2241-2258页

作者： Goodman, John L. Odyssey Space Res LLC Houston TX 77058 USA

The mission requirements of crewed space vehicles flown by NASA drove the development of new techniques for burn targeting and guidance for ascent, rendezvous, lunar landing, and deorbit. The Saturn V Iterative Guidance Mode used an approximately optimal steering law that could adjust the trajectory to compensate for performance issues while meeting trajectory constraints at propulsion-system cutoff. Hypersurface targeting provided constraints for the Saturn V Trans-Lunar Injection burn. E Guidance, in quadratic- and linear-acceleration forms, provided closed-loop descent and ascent guidance for the Apollo Lunar Module. The Space Shuttle's Powered Explicit Guidance algorithm was more flexible than Apollo-era guidance algorithms, and it had better convergence characteristics, enabling it to support challenging abort profiles. The Space Launch System has built on the heritage of the Space Shuttle's guidance algorithm. The Orion spacecraft's Two-Level Targeter enables onboard, autonomous targeting of an entire mission profile subject to three-body dynamics in cislunar space. Orion uses Shuttle-derived guidance as the basis for an orbit-guidance package with more capability than guidance algorithms of previous vehicles.

关键词： NASA Guidance and Navigational algorithms Exploration Upper Stage Clohessy Wiltshire Equation Sample Return Missions Satellite Launch Vehicle Crewed Spacecrafts Space Exploration and Technology targeting algorithms Optimization Algorithm

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Onboard Autonomous targeting for the Trans-Earth Phase of Orion

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2010年第3期33卷 943-956页

作者： Marchand, Belinda G. Weeks, Michael W. Smith, Chad W. Scarritt, Sara Univ Texas Austin Austin TX 78712 USA NASA Lyndon B Johnson Space Ctr Aerosci & Flight Mech Div Code EG6 Houston TX 77058 USA

The present investigation focuses on one aspect of the autonomous targeting process used onboard during the Orion trans-Earth injection phase, specifically, a fast and robust algorithm that identifies a feasible return trajectory, one that meets the entry constraints without exceeding the fuel available. Unlike earlier Apollo missions, Orion seeks to land near the polar regions of the moon. Thus, a substantial plane change maneuver is required before returning to Earth. To reduce the fuel expenditure associated with this plane change, a three-maneuver sequence is employed during the return phase. An autonomous onboard targeting process for precision entry (one that incorporates multiple coordinated trans-Earth maneuvers) is sought in the event of loss of communication with the ground. The latter scenario presents a very unique challenge: one never before required of any Apollo vehicle. The Apollo missions also benefited from flexible entry requirements in contrast to Orion. Precision targeting in multibody regimes has only been previously demonstrated in unmanned sample return missions such as Genesis. The formulation presented here ensures that the entry constraints are met without violating the available fuel budget.

关键词： Earth Trans Earth Injection targeting algorithms Moon Apollo Missions Sample Return Missions Flight Path Angle Earth Centered Inertial Sequential Quadratic Programming FORTRAN

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targeting Ballistic Lunar Capture Trajectories Using Periodic Orbits

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2011年第3期34卷 893-902页

作者： Griesemer, Paul Ricord Ocampo, Cesar Cooley, D. S. Univ Texas Austin Dept Aerosp Engn & Engn Mech Austin TX 78712 USA NASA Goddard Space Flight Ctr Nav & Miss Design Branch Greenbelt MD 20771 USA

A particular periodic orbit in the Earth sun circular restricted three-body problem is shown to have the characteristics needed for a ballistic lunar capture transfer. An injection from a circular parking orbit into the periodic orbit serves as an initial guess for a nonlinear targeting algorithm. By targeting appropriate parameters incrementally in increasingly complicated force models and using precise derivatives calculated from the state transition matrix, a reliable algorithm is produced. Ballistic lunar capture trajectories in restricted four-body systems are produced in a systematic way.

关键词： Moon Earth Elliptical Orbit Libration Points Ecliptic targeting algorithms Low Energy Transfer Perigee Sequential Quadratic Programming Orbital Elements

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Spacecraft Deorbit Point targeting Using Aerodynamic Drag

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2017年第10期40卷 2646-2652页

作者： Omar, Sanny R. Bevilacqua, Riccardo Guglielmo, David Fineberg, Laurence Treptow, Justin Clark, Scott Johnson, Yusef Univ Florida ADAMUS Lab Dept Mech & Aerosp Engn 939 Sweetwater DrMAE-A Room 211 Gainesville FL 32611 USA Univ Florida ADAMUS Lab Dept Mech & Aerosp Engn 939 Sweetwater DrMAE-A Room 308 Gainesville FL 32611 USA Univ Florida ADAMUS Lab Dept Mech & Aerosp Engn 939 Sweetwater Dr Gainesville FL 32611 USA NASA Kennedy Space Ctr Launch Serv Program Kennedy Space Ctr FL 32899 USA Ai Solut Inc NASA Kennedy Space Ctr Titusville FL 32899 USA

Presented as Paper 2017-1268 at the AIAA Guidance, Navigation, and Control Conference, Grapevine, TX, 9-13 January 2017

Presented as Paper 2017-1268 at the AIAA Guidance, Navigation, and Control Conference, Grapevine, TX, 9-13 January 2017

关键词： Spacecrafts Aerodynamic Drag Satellites Ballistic Coefficient targeting algorithms Earth Earth Centered Inertial Cumulative Distribution Function Right Ascension MATLAB

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Orbital targeting Using Reduced Eccentric Anomaly Low-Thrust Coefficients

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2011年第3期34卷 820-831页

作者： Hudson, Jennifer S. Scheeres, Daniel J. Univ Michigan Dept Aerosp Engn Ann Arbor MI 48019 USA Univ Colorado Dept Aerosp Engn Sci Colorado Ctr Astrodynam Res Boulder CO 80309 USA

A method to evaluate the trajectory dynamics of low-thrust spacecraft is applied to targeting and optimal control problems. Averaged variational equations for the osculating orbital elements are used to estimate a spacecraft trajectory over many spiral orbits. Fourteen Fourier coefficients of the thrust acceleration vector represent the fundamental trajectory dynamics. Spacecraft targeting problems are solved using the averaged variational equations and a general cost function represented as a Fourier series. The resulting fuel costs and dynamic fidelity of the targeting solutions are evaluated. The goal of the method is not precise targeting, but easy reconstruction of the basic elements of the thrusting trajectory and control law.

关键词： Thrust Eccentric Anomaly Orbital Elements Space Situational Awareness Boundary Value Problems Solar Electric Propulsion targeting algorithms Spacecraft Trajectories Taylor Series Thrust Vectoring

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Navigation Solution to Maneuver Spacecraft Relative to Spheres Centered on Cooperative Satellites

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JOURNAL OF SPACECRAFT AND ROCKETS 2015年第3期52卷 754-765页

作者： Leigh, Abraham M. Black, Jonathan T. US Air Force Inst Technol Dept Aeronaut & Astronaut Engn Wright Patterson AFB OH 45433 USA

A differential correction algorithm is presented to deliver an impulsive maneuver to a satellite to place it within the overlapping spheres, with user-defined radii, centered around multiple cooperative satellites within a constrained time. The differential correction algorithm develops and utilizes the state transition matrix along with the equations of motion and multiple satellites' state information to determine the optimum trajectory to achieve the desired results. The results from the algorithm are presented for prograde orbits. The results are presented in order to allow for mission design trade-offs, including satellite separation distance and required V based on the satellites' initial inclinations. The results also provide empirical methods to determine the miss distance and optimum V solutions for the provided problem. This work ultimately provides the framework for applying the algorithm to a unique user defined responsive space scenario.

关键词： Satellites Spacecrafts Orbital Eccentricity Earth Lambert Algorithm Highly Elliptical Orbit Ballistic Coefficient Orbit Determination targeting algorithms Terminal Guidance

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Minimum-Landing-Error Powered-Descent Guidance for Mars Landing Using Convex Optimization

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2010年第4期33卷 1161-1171页

作者： Blackmore, Lars Acikmese, Behcet Scharf, Daniel P. CALTECH Jet Prop Lab Guidance & Control Anal Grp Pasadena CA 91109 USA

To increase the science return of future missions to Mars and to enable sample return missions, the accuracy with which a lander can be delivered to the Martian surface must be improved by orders of magnitude. The prior work developed a convex-optimization-based minimum-fuel powered-descent guidance algorithm. In this paper, this convex-optimization-based approach is extended to handle the case when no feasible trajectory to the target exists. In this case, the objective is to generate the minimum-landing-error trajectory, which is the trajectory that minimizes the distance to the prescribed target while using the available fuel optimally. This problem is inherently a nonconvex optimal control problem due to a nonzero lower bound on the magnitude of the feasible thrust vector. It is first proven that an optimal solution of a convex relaxation of the problem is also optimal for the original nonconvex problem, which is referred to as a lossless convexification of the original nonconvex problem. Then it is shown that the minimum-landing-error trajectory generation problem can be posed as a convex optimization problem and solved to global optimality with known bounds on convergence time. This makes the approach amenable to onboard implementation for real-time applications.

关键词： Powered Descent Guidance Convex Optimization Mars Science Laboratory Optimal Control Mars Exploration Rover Jet Propulsion Laboratory targeting algorithms Nonlinear Systems Mars Pathfinder Thrust Vectoring

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Analytical Lunar Descent Guidance Algorithm

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2009年第3期32卷 915-926页

作者： Chomel, Christina T. Bishop, Robert H. Univ Texas Austin Appl Res Labs Adv Technol Lab Austin TX 78713 USA Univ Texas Austin Dept Aerosp Engn & Engn Mech Univ Stn 1 Austin TX 78712 USA

The desire to return humans to the moon has motivated the evaluation of existing lunar descent guidance algorithms and the development of new ideas. This paper outlines a design for a targeting algorithm that quickly and reliably generates a two-dimensional reference trajectory and a real-time three-dimensional guidance algorithm that can use this or any reference trajectory as its basis. The targeting algorithm is analytical in nature and allows for updates to the reference in real time. Monte Carlo analysis demonstrates the reliable performance of the algorithms in both a general sense and for retargeting once the lunar descent has been initiated.

关键词： Guidance algorithms targeting algorithms Moon Monte Carlo Analysis Flight Path Angle Gravity Turn Thrust Vectoring Johnson Space Center Proportional Control Descent Guidance

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ESTIMATION OF THE TIME-TO-GO PARAMETER FOR AIR-TO-AIR MISSILES

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 1985年第2期8卷 262-266页

作者： LEE, GKF Colorado State Univ. Fort Collins CO USA

The design of optimal controllers for air-to-air missiles has been of much recent interest. Many control laws require a good estimate of the time-to-go (tgo) before intercept. This paper presents an algorithm for estimating time-to-go using the current relative position and velocity characteristics of the missile and target. By constructing and objective function representing terminal miss distance, one is able to obtain an iterative scheme for estimating tgo. Simulation studies using a smart target algorithm and this methodology have yielded good results that show improvement over conventional algorithms for estimating tgo. A comparative investigation of the methodology developed in this paper with other algorithm is presented.

关键词： Air to Air Missiles targeting algorithms Guidance Laws Transversality Condition Proportional Navigation Missile Body Air Force Base Aerodynamic Properties Closed Loop System Flight Maneuvers

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Event Triggers in Linear Covariance Analysis with Applications to Orbital Rendezvous

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2009年第1期32卷 102-111页

作者： Geller, David K. Ben Rose, M. Woffinden, David C. Utah State Univ Logan UT 84322 USA

The analytic expressions for a generic event trigger for a linear covariance analysis are developed. The event triggers are applied to the analysis of a common orbital rendezvous problem. Some unexpected results are observed and a flaw in an equation that is commonly used in standard linear covariance analysis applications is uncovered and remedied. A quantitative analysis of the effect of elevation-angle triggers for various rendezvous scenarios is presented.

关键词： Earth Orbital Rendezvous Orbital Transfer Monte Carlo Simulation Flight Computer Resident Space Object LIDAR targeting algorithms White Noise Navigation Algorithm Dirac Delta Function

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