Mission analysis of the D'Arrest '76 opportunity
Comet de Arrest 1976 opportunity mission analysis, considering flyby and rendezvous trajectories and guidance problems
Roberts, P. H., Jr.↗
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The Halley rendezvous via solar sailing mission description
The possibility of using solar sails to achieve a rendezvous with Halley's Comet is explored. Rendezvous trajectories are calculated.
Friedman, L. D.↗
Rendezvous guidance
Viewgraphs on rendezvous guidance are presented. Topics covered include: rendezvous guidance; Lambert Targeting; Linear Quadratic Targeting; unmanned vehicles; rendezvous trajectory; and fuzzy logic control.
Mcguire, Andrew T.↗
Interactive orbital proximity operations planning system instruction and training guide
This guide instructs users in the operation of a Proximity Operations Planning System. This system uses an interactive graphical method for planning fuel-efficient rendezvous trajectories in the multi-spacecraft environment of the space station and allows the operator to compose a multi-burn transfer trajectory between orbit initial chaser and target trajectories. The available task time (window) of the mission is predetermined and the maneuver is subject to various operational constraints, such as departure, arrival, spatial, plume impingement, and en route passage constraints. The maneuvers are described in terms of the relative motion experienced in a space station centered coordinate system. Both in-orbital plane as well as out-of-orbital plane maneuvering is considered. A number of visual optimization aids are used for assisting the operator in reaching fuel-efficient solutions. These optimization aids are based on the Primer Vector theory. The visual feedback of trajectory shapes, operational constraints, and optimization functions, provided by user-transparent and continuously active background computations, allows the operator to make fast, iterative design changes that rapidly converge to fuel-efficient solutions. The planning tool is an example of operator-assisted optimization of nonlinear cost functions.
Grunwald, Arthur J.↗
Guidance, navigation, and control study for a solar electric propulsion spacecraft
A preliminary investigation of a lunar-comet rendezvous mission using a solar electric propulsion (SEP) spacecraft was performed in two phases.The first phase involved exploration of the moon and the second involved rendezvous with a comet. The initial phase began with a chemical propulsion translunar injection and chemical insertion into a lunar orbit, followed by a low thrust SEP transfer to a circular, polar, low-lunar orbit. After collecting scientific data at the moon, the SEP spacecraft performed a spiral lunar escape maneuver to begin the interplanetary leg of the mission. After escape from the Earth-moon system, the SEP spacecraft maneuvered in interplanetary space and performed a rendezvous with a comet.The immediate goal of this study was to demonstrate the feasibility of using a low-thrust SEP spacecraft for orbit transfer to both the moon and a comet. Another primary goal was to develop a computer optimization code which would be robust enough to obtain minimum-fuel rendezvous trajectories for a wide range of comets.
Kluever, Craig A.↗
Interactive method for planning constrained, fuel-optimal orbital proximity operations
An interactive graphical method for planning fuel-efficient rendezvous trajectories in the multi-spacecraft environment of the space station is presented. The method allows the operator to compose a multi-burn transfer trajectory between arbitrary initial chaser and target trajectories. The available task time of the mission is limited and the maneuver is subject to various operational constraints, such as departure, arrival, plume impingement and spatial constraints. The maneuvers are described in terms of the relataive motion experienced in a Space-Station centered coordinate system. The optimization method is based on the primer vector and its extension to non-optimal trajectories. The visual feedback of trajectory shapes, operational constraints, and optimization functions, provided by user-transparaent and continuously active background computations, allows the operator to make fast, iterative design changes which rapidly converge to fuel-efficient solutions. The optimization functions are presented. A variety of simple design examples has been presented to demonstrate the usefulness of the method. In many cases the addition of a properly positioned intermediate waypoint resulted in fuel savings of up to 30%. Furthermore, due to the counter-intuitive character of the optimization functions, most fuel-optimal solutions could not have been found without the aid of the optimization tools. Operating the system was found to be very easy, and did not require any previous in-depth knowledge of orbital dynamics or trajectory. The planning tool is an example of operator assisted optimization of nonlinear cost-functions.
Abramovitz, Adrian↗
Methodology for Prototyping Increased Levels of Automation for Spacecraft Rendezvous Functions
The Crew Exploration Vehicle necessitates higher levels of automation than previous NASA vehicles, due to program requirements for automation, including Automated Rendezvous and Docking. Studies of spacecraft development often point to the locus of decision-making authority between humans and computers (i.e. automation) as a prime driver for cost, safety, and mission success. Therefore, a critical component in the Crew Exploration Vehicle development is the determination of the correct level of automation. To identify the appropriate levels of automation and autonomy to design into a human space flight vehicle, NASA has created the Function-specific Level of Autonomy and Automation Tool. This paper develops a methodology for prototyping increased levels of automation for spacecraft rendezvous functions. This methodology is used to evaluate the accuracy of the Function-specific Level of Autonomy and Automation Tool specified levels of automation, via prototyping. Spacecraft rendezvous planning tasks are selected and then prototyped in Matlab using Fuzzy Logic techniques and existing Space Shuttle rendezvous trajectory algorithms.
Hart, Jeremy J.↗
Cost-Effective Icy Bodies Exploration using Small Satellite Missions
It has long been known that Saturn's moon Enceladus is expelling water-rich plumes into space, providing passing spacecraft with a window into what is hidden underneath its frozen crust. Recent discoveries indicate that similar events could also occur on other bodies in the solar system, such as Jupiter's moon Europa and the dwarf planet Ceres in the asteroid belt. These plumes provide a possible giant leap forward in the search for organics and assessing habitability beyond Earth, stepping stones toward the long-term goal of finding extraterrestrial life. The United States Congress recently requested mission designs to Europa, to fit within a cost cap of $1B, much less than previous mission designs' estimates. Here, innovative cost-effective small spacecraft designs for the deep-space exploration of these icy worlds, using new and emerging enabling technologies, and how to explore the outer solar system on a budget below the cost horizon of a flagship mission, are investigated. Science requirements, instruments selection, rendezvous trajectories, and spacecraft designs are some topics detailed. The mission concepts revolve around a comparably small-sized and low-cost Plume Chaser spacecraft, instrumented to characterize the vapor constituents encountered on its trajectory. In the event that a plume is not encountered, an ejecta plume can be artificially created by a companion spacecraft, the Plume Maker, on the target body at a location timed with the passage of the Plume Chaser spacecraft. Especially in the case of Ceres, such a mission could be a great complimentary mission to Dawn, as well as a possible future Europa Clipper mission. The comparably small volume of the spacecraft enables a launch to GTO as a secondary payload, providing multiple launch opportunities per year. Plume Maker's design is nearly identical to the Plume Chaser, and fits within the constraints for a secondary payload launch. The cost-effectiveness of small spacecraft missions enables the exploration of multiple solar system bodies in reasonable timeframes despite budgetary constraints, with only minor adaptations. The work presented here is a summary of concepts targeting icy bodies, such as Europa and Ceres, which have been developed over the last year at NASA Ames Research Center's Mission Design Division. The platforms detailed in this work are also applicable to the cost-effective exploration of many other small icy bodies in the solar system.
Mission Design↗
Angles-Only Robust Trajectory Optimization for NRHO Rendezvous
This study demonstrates a robust trajectory optimization approach for rendezvous and proximity operations with angles-only navigation measurements. Often, sensors that directly measure relative range and velocity require communication or coordination between the chaser and target vehicle and can have limiting pointing accuracy, mass, or power requirements compared to angle measurement sensors. Thus, the capability to perform a rendezvous with only angle measurements can be advantageous for vehicle design and to improve robustness to failures. However, the well studied limitation of angles-only navigation in measuring range results in large uncertainties in the navigation system that must be reduced with chaser vehicle thrust maneuvers to induce observability in range for the navigation filter. This analysis presents a trajectory optimization problem for a lunar ascent rendezvous during a crewed lunar mission in a Near-Rectilinear Halo Orbit (NRHO) that is limited to only angle measurements. The objective of this study is to show that an angles-only rendezvous is feasible in an NRHO and to present the sensitivity to an assortment of constraints generated from a systematic optimization process using linear covariance analysis and particle swarm optimization. Linearized NRHO dynamics and linearized relative targeting are applied to use linear covariance analysis to determine the expected delta-v and trajectory dispersions due to initial state uncertainty, sensor errors, maneuver execution errors, and unmodeled dynamics. The delta-v and trajectory dispersions are passed into a particle swarm optimization algorithm to find the optimized maneuver profile that minimizes fuel use while satisfying constraints such as free drift and underburn to 3-sigma certainty. The trajectory constraints including time available, desired final uncertainty, and initial uncertainty are varied to ascertain sensitivity and desirable engineering trades.
Linear Covariance Analysis↗
Near Earth Asteroid Rendezvous Orbit Phase Trajectory Design
Trajectory design of the orbit phase of the Near Earth Asteroid Rendezvous (NEAR) mission involves procedures that depart significantly from those used for previous missions.
Missions↗
Robust Trajectory Design for Rendezvous in a Near Rectilinear Halo Orbit
Future NASA Artemis missions will require complex docking plans between the Orion capsule and Lunar Gateway that meet predetermined safety constraints while minimizing fuel usage and state uncertainty at rendezvous. In this paper, linear covariance analysis is applied to a first order relative form of Near-Rectilinear Halo Orbit dynamics to determine the nominal trajectories and state dispersions associated with various maneuver profiles in the Sun-referenced Local Vertical Local Horizontal reference frame of the lunar Gateway. These maneuver profiles are optimized using a particle swarm optimizer and direct search algorithm to find trajectories that satisfy approach corridor, free drift, velocity magnitude, under-burn, and maneuver transfer time safety constraints to 3-sigma certainty.
Linear Covariance Analysis↗
Trajectory design for the Comet Rendezvous Asteroid Flyby 1995-1996 opportunities
Recent reprogramming of NASA's Comet Rendezvous Asteroid Flyby (CRAF) mission has motivated an examination of the comet rendezvous opportunities available with launch dates in 1995 and 1996. Various interplanetary transfer modes have been investigated including direct, Delta V-earth-gravity- assist and Venus/earth gravity assist trajectories. Optimal trajectories have been generated and their performance assessed for a spacecraft using a chemical propulsion system and the upgraded Titan/Centaur launch vehicle. Of the viable options available, a new CRAF baseline mission has been selected. One important result of the present study is that while there are comet rendezvous trajectories with launch dates throughout the next decade, 1997 marks the beginning of a hiatus of opportunities with short flight times to the comet targets which promise the greatest potential science return.
Myers, Mark R.↗
Encounter trajectory design for SEP rendezvous with low mass celestial bodies.
This paper shows the considerations which precede and affect encounter trajectory design and the impact of this design on propulsion system performance requirements and interplanetary trajectory design. Representative missions selected to illustrate the problems and characteristics of encounter trajectory design are a Mercury Orbiter, Eros Rendezvous, Encke Rendezvous, and Ceres Orbiter. The paper shows that for SEP missions, particularly for low mass celestial bodies, the encounter trajectory may be freely specified to a large extent to satisfy mission goals rather than being dictated by a rigid interplanetary trajectory design.
Driver, J. M.↗
A dual phase plane approach to the two-finite- burns minimum fuel rendezvous problem.
Two-finite burns minimum fuel rendezvous problem using dual phase plane description, determining control law
VARIATION METHOD↗
Application of an impulsive trajectory optimization method to the comet rendezvous problem
Multiple impulsive spacecraft trajectory optimization technique application to comet rendezvous problem, using computer program
Waters, J. I.↗
Robust Trajectory Optimization for NRHO Rendezvous Using SPICE Kernel Relative Motion
In this paper, robust optimization is performed on trajectory correction maneuvers during the lunar lander return phase of an Artemis mission, treating the trajectory from one hour after low lunar orbit departure to arrival in the vicinity of the lunar Gateway as a relative motion problem. To enable rapid stochastic optimization techniques requiring many candidate trajectories, SPICE kernel relative motion as implemented by the Quadratic Interpolated State Transition (QIST) system is used as the underlying dynamics propagation. The optimization is performed with a genetic optimizer using linear covariance (LinCov) software in a simplified operational context, taking into account the availability of navigation sensors with varying measurement models, ranges, and accuracies. No numerical integration is used, since the relative motion around Gateway is fully characterized with the a priori computation of the QIST coefficients. Maneuver placements are computed to optimize the minimum 3σ delta-v of the trajectory, the position dispersion at a target point, and a convex combination of these two metrics. An order of magnitude runtime improvement is provided over legacy methods with less than 10% error introduced. All QIST results are shown to be in-family with legacy methods. The tradespace for optimal delta-v design is found to range from 77.0 to 93.9 m/s, while the range of optimal dispersion is between 1.4 and 11.7 km.
Relative Motion↗
On energy-optimal trajectories in three dimensions for the terminal phase of satellite rendezvous
Energy-optimal trajectories for terminal phase of satellite rendezvous
TRAJECTORY↗
A Study of the Positions and Velocities of a Space Station and a Ferry Vehicle During Rendezvous and Return
A study is made of the families of non-thrusting ascent trajectories during rendezvous with an orbiting space station and the descent trajectories to the earth's atmosphere. Equations of motion are derived and results are shown for two typical orbits of the station (one circular and one elliptic orbit}. Boundaries of launch (at time of booster burnout) and rendezvous conditions are given and the effects of delays in launch time are discussed.
Eggleston, John M.↗