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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Tracking and data system support for the Mariner Mars 1971 mission. Volume 2: First trajectory correction maneuver through orbit insertion

The Deep Space Tracking and Data System activities in support of the Mariner Mars 1971 project from the first trajectory correction maneuver on 4 June 1971 through cruise and orbit insertion on 14 November 1971 are presented. Changes and updates to the TDS requirements and to the plan and configuration plus detailed information on the TDS flight support performance evaluation and the preorbital testing and training are included. With the loss of Mariner 8 at launch, a few changes to the Mariner Mars 1971 requirements, plan, and configuration were necessitated. Mariner 9 is now assuming the former mission plan of Mariner 8, including the TV mapping cycles and a 12-hr orbital period. A second trajectory correction maneuver was not required because of the accuracy of the first maneuver. All testing and training for orbital operations were completed satisfactorily and on schedule. The orbit insertion was accomplished with excellent results.

Textor, G. P.↗
NASA NTRS1988

Galileo Earth-Venus trajectory correction maneuver design

As a result of delays in the Space Shuttle program and a re-evaluation of safety concerns which resulted in cancellation of the Shuttle Centaur Upper Stage program, the Galileo mission to Jupiter has undergone a substantial redesign effort. Use of the Inertial Upper Stage (IUS) for injection has necessitated a complicated Venus-Earth-Earth gravity assist trajectory in order to reach Jupiter. One flyby of Venus and two flybys of Earth are required to compensate for the lower injection energy available from the IUS. As the spacecraft was not originally designed for the environment within 1 AU of the Sun, modifications to the spacecraft and operating procedures have been developed. This paper describes some of these changes and how they have affected the design and implementation of trajectory correction maneuvers on the Earth-Venus leg of the mission. A strategy of biasing the IUS target at injection to control the statistical velocity distribution of a subsequent trajectory correction maneuver is discussed.

Wilson, M. G.↗
NASA NTRS1982

Statistics of Delta v magnitude for a trajectory correction maneuver containing deterministic and random components

A number of interplanetary missions now being planned involve placing deterministic maneuvers along the flight path to alter the trajectory. Lee and Boain (1973) examined the statistics of trajectory correction maneuver (TCM) magnitude with no deterministic ('bias') component. The Delta v vector magnitude statistics were generated for several values of random Delta v standard deviations using expansions in terms of infinite hypergeometric series. The present investigation uses a different technique (Monte Carlo simulation) to generate Delta v magnitude statistics for a wider selection of random Delta v standard deviations and also extends the analysis to the case of nonzero deterministic Delta v's. These Delta v magnitude statistics are plotted parametrically. The plots are useful in assisting the analyst in quickly answering questions about the statistics of Delta v magnitude for single TCM's consisting of both a deterministic and a random component. The plots provide quick insight into the nature of the Delta v magnitude distribution for the TCM.

Bollman, W. E.↗
NASA NTRS2021

Mars 2020 Trajectory Correction Maneuver Design

The Mars 2020 mission launched on 30 July 2020 and arrived at Mars on 18 February 2021. Delivering the mission’s Perseverance rover to Jezero Crater required adjusting the postlaunch trajectory to remove the launch injection bias and to target the atmospheric entry conditions, while also satisfying requirements on the propellant usage and non-nominal impact probability at Mars. The Mars 2020 maneuver design team achieved these goals by designing and executing three propulsive maneuvers in flight. The accurate on-board execution of those maneuvers delivered the spacecraft into the Martian atmosphere in a state that allowed for a successful entry, descent, and landing on the surface. This paper details the maneuver design process and describes the design and execution of the three in-flight propulsive maneuvers.

Kruizinga, Gerhard↗
NASA NTRS1972

Tracking and data system support for the Mariner Mars 1971 mission. Prelaunch phase through first trajectory correction maneuver, volume 1

The DSN command system provided the capability to enter commands in a computer at the deep space stations for transmission to the spacecraft. The high-rate telemetry system operated at 16,200 bits/sec. This system will permit return to DSS 14 of full-resolution television pictures from the spacecraft tape recorder, plus the other science experiment data, during the two playback periods of each Goldstone pass planned for each corresponding orbit. Other features included 4800 bits/sec modem high-speed data lines from all deep space stations to Space Flight Operations Facility (SFOF) and the Goddard Space Flight Center, as well as 50,000 bits/sec wideband data lines from DSS 14 to the SFOF, thus providing the capability for data flow of two 16,200 bits/sec high-rate telemetry data streams in real time. The TDS performed prelaunch training and testing and provided support for the Mariner Mars 1971/Mission Operations System training and testing. The facilities of the ETR, DSS 71, and stations of the MSFN provided flight support coverage at launch and during the near-earth phase. The DSSs 12, 14, 41, and 51 of the DSN provided the deep space phase support from 30 May 1971 through 4 June 1971.

Laeser, R. P.↗
NASA NTRS1984

Design and analysis techniques for trajectory correction maneuvers

Selecting a spacecraft maneuver strategy usually involves tradeoffs of many competing factors. This paper describes such a tradeoff process. The method uses parametric data to cope with targeting requirements, thruster configuration and performance, hardware constraints, operational considerations, propellant optimization demands, and expected execution and orbit determination errors, while remaining flexible to react to new conditions.

Hintz, G. R.↗
NASA NTRS1986

Precision maneuver determination for trajectory corrections

The interplanetary navigation of a spacecraft requires precise calculation of trajectory correction maneuver parameters. Velocity increments must be found which make the required corrections, minimize propellant consumption, and do not violate hardware design constraints or scientific objectives. This process is illustrated by its application to the Galileo Mission. The paper describes inflight analysis techniques, mathematical modeling, and computational algorithms used in the precision maneuver calculation scenario.

Hintz, G. R.↗
NASA NTRS1986

High-accuracy computation of Delta V magnitude probability densities - Preliminary remarks

This paper describes an algorithm for the high accuracy computation of some statistical quantities of the magnitude of a random trajectory correction maneuver (TCM). The trajectory correction velocity increment Delta V is assumed to be a three-component random vector with each component being a normally distributed random scalar having a possibly nonzero mean. Knowledge of the statitiscal properties of the magnitude of a random TCM is important in the planning and execution of maneuver strategies for deep-space missions such as Galileo. The current algorithm involves the numerical integration of a set of differential equations. This approach allows the computation of density functions for specific Delta V magnitude distributions to high accuracy without first having to generate large numbers of random samples. Possible applications of the algorithm to maneuver planning, planetary quarantine evaluation, and guidance success probability calculations are described.

Chadwick, C.↗
NASA NTRS2019

Charting a course to the sun: flight path control for Parker Solar Probe

The successful launch of the Parker Solar Probe (PSP) on August 12, 2018 with a Delta IV rocket and Star-48BV third stage has placed the spacecraft on a 7-year trajectory to study the Sun. The goals of PSP are to better characterize our solar environment and advance our understanding of the Sun at 9.86 Rs. A total of 42 trajectory correction maneuvers are planned. This paper documents trajectory correction maneuver analysis performed just prior to launch until just past the first solar encounter. The pre-launch analysis culminated in two final design cycles which analyzed 24 reference trajectories.

Ryne, Mark↗
NASA NTRS1975

Mariner Jupiter/Saturn 1977 navigation strategy

This paper presents selected navigational aspects of the Mariner Jupiter/Saturn 1977 mission. Emphasis is given to the trajectory correction strategy. In addition to science return and propellant costs, specific considerations affecting the strategy include a large number of candidate trajectories, relatively close satellite flybys, use of trajectory correction maneuvers for trajectory shaping, and the requirement for an optical data type to achieve precision satellite encounters. Finally, important error sources and mission constraints are identified, and their impact on the strategy is discussed.

Mckinley, E. L.↗
NASA NTRS1992

Ulysses mission operations

The Ulysses mission is described in terms of in-Shuttle operations, initial in-orbit operations, routine operations, operational organization, and data gathering and production. The configuration of the Ulysses payload is illustrated, and the flight to orbit is described including a three-hour on-orbit checkout. The first contact was reported at the Deep Space Network station followed by an adjustment of the spacecraft solar-aspect angle and the acquisition of ranging and Doppler data. In-orbit operations include the earth acquisition maneuver, a trajectory correction maneuver, and a payload switch. Continuous data gathering is discussed with reference to the Jupiter encounter and the first and second oppositions and conjunctions. The data-gathering components comprise ground stations, a data-processing computer, and a data-records system. Data production is performed in an off-line mode that does not interfere with the real-time operations.

Beech, P.↗
NASA NTRSDate not supplied

Electric Sail Design Sensitivities

Electric Sails (E-Sails) are a promising propulsion technology that seek to enable high characteristic propellentless acceleration for spacecraft to reach distant and/or difficult to reach orbits such as rapid transit to heliopause. The E-Sail system exchanges momentum by using positively charged electrostatic tethers to repel solar wind photons to push it through space. This concept was first theorized by Pekka Janhunen in 2004 with further developments occurring including a NASA Innovative Advanced Concepts Phase 1 and Phase 2 hosted out of NASA’s Marshall Space Flight Center (MSFC). These developments led to further maturation of the overall system, and this paper is designed to help identify different design sensitivities of an integrated Electric Sail system. To approach this, a small team at MSFC took an in-house developed three degrees of freedom (3DoF) simulation tool and a trajectory modelling tool to look at different design parameters and to determine a potential ideal E-Sail configuration. An E-Sail system has different design architectures including a barbell design, hub and spoke design, and a potential hybrid solution. The barbell design features two equally massed satellites that spin around a central point in the tether system. The hub and spoke design features a large central spacecraft with small end spacecraft to aid in formation control of the overall E-Sail system. Leveraging elements from both the hub and spoke, and barbell design, a hybrid option exists where one could have a larger central mass and one or two tethers extended to a smaller end mass [2]. In summary, the hub and spoke design is the ideal configuration for E-Sail with tethers spanning kilometers to achieve the designed design characteristic acceleration of at least one mm/s2 with this architecture being the focus of this study. Different key design parameters were varied as part of this study. These parameters include the total number of tethers, the length of the tethers, spin rate, relative spacecraft mass, tether voltage, inertia per tether, and impact of slew rate changes. These results started from an internal MSFC technology demonstration mission design and then the parameters were varied with engineering judgement to ensure that the results were consistent with expected results. These were varied in the MSFC 3DoF tool and compared to baseline results to generate a candidate ideal E-Sail design system. These results were then used to help inform the mission analysis design for the system. The 3DoF trajectory optimization was performed in the Astrodynamics and Space Science Enabling Toolbox (ASSET) and utilized collocation optimization of control vector pointing with the objective function being transfer duration. A representative model of the E-Sail force was derived as a function of the Sun-pointing vector direction. This model was initialized with a zero Sun Incidence Angle (SIA) trajectory for the initial guess generation, then with a portion of the mission optimized SIA time history for fast outbound transfers. After the trajectory is hyperbolic relative to a heliocentric frame, the SIA is maintained at 0 degrees except for any needed trajectory correction maneuvers. This trajectory optimization looked at different characteristic acceleration values to analyze the impact it had on mission performance. As a result of this study, the team can best inform mission designers and technology development efforts to mature the E-Sail system. This study will allow mission designers to have defined rules-of-thumb to design an E-Sail system to meet desired mission needs as well as providing several sample mission profiles. These design drivers will inform maturation efforts on which design requirements to consider for an integrated E-Sail design.

E-Sail↗
NASA NTRS2004

Mars Exploration Rovers Propulsive Maneuver Design

The Mars Exploration Rovers Spirit and Opportunity successfully landed respectively at Gusev Crater and Meridiani Planum in January 2004. The rovers are essentially robotic geologists, sent on a mission to search for evidence in the rocks and soil pertaining to the historical presence of water and the ability to possibly sustain life. In order to conduct NASA's 'follow the water' strategy on opposite sides of the planet Mars, an interplanetary journey of over 300 million miles culminated with historic navigation precision. Rigorous trajectory targeting and control was necessary to achieve the atmospheric entry requirements for the selected landing sites. The propulsive maneuver design challenge was to meet or exceed these requirements while preserving the necessary design margin to accommodate additional project concerns. Landing site flexibility was maintained for both missions after launch, and even after the first trajectory correction maneuver for Spirit. The final targeting strategy was modified to improve delivery performance and reduce risk after revealing constraining trajectory control characteristics. Flight results are examined and summarized for the six trajectory correction maneuvers that were planned for each mission.

trajectory↗
NASA NTRS1994

(abstract) Maneuver Design for Galileo Asteroid Flybys

This paper will address the trajectory correction maneuvers (TCMs) which enabled the successful completion of the circuitous VEEGA trajectory while providing for the first ever encounters with main-belt asteroids.

Galileo trajectory correction maneuvers Jupiter as↗
NASA NTRS2025

Mission-Maps For Outbound Cislunar Transfer Trajectories

This study quantifies the robustness and sensitivity of an outbound cislunar trajectory for a lunar lander in the form of mission-maps, or topological maps that allows either a computer program or mission designer to intuitively optimize the placement of critical outbound correction burns from the derived sensitivity data. The non-linear multi-body dynamics are applied to generate an outbound cislunar reference profile used by a linear covariance analysis (LinCov) tool to compute the expected Δv and trajectory dispersions due to the initial state uncertainty, sensor errors, maneuver execution errors, and disturbance accelerations along the outbound cislunar profile. The rapid performance analysis capabilities of LinCov are complimented with parallel processing techniques to evaluate hundreds and thousands of different translational burn locations, placements, and targeting constraints to identify the combination that minimizes the total Δv usage (nominal plus 3σ Δv) and trajectory dispersions at lunar orbit insertion. This study utilizes a generalized reference targeting algorithm to quickly assess the integrated closed-loop GN&C system performance due to different targeting configurations and constraints. The resulting mission maps provide an intuitive insight to ascertain each trajectory correction maneuver’s (TCM) sensitivity to different burn times along an outbound cislunar trajectory and quickly identify desirable engineering tradeoffs when performing analysis on the number and placement of these burns that nominally zero. Multiple mission maps are generated for a variety of different performance parameters that allow engineers to visually identify optimal solutions for trajectory correction maneuver placements, the number of correction burns, and the targeting constraints for each burn.

GN&C↗
NASA NTRS2015

OSIRIS-REx, Returning the Asteroid Sample

This paper addresses the technical aspects of the sample return system for the upcoming Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer (OSIRIS-REx) asteroid sample return mission. The overall mission design and current implementation are presented as an overview to establish a context for the technical description of the reentry and landing segment of the mission.The prime objective of the OSIRIS-REx mission is to sample a primitive, carbonaceous asteroid and to return that sample to Earth in pristine condition for detailed laboratory analysis. Targeting the near-Earth asteroid Bennu, the mission launches in September 2016 with an Earth reentry date of September 24, 2023.OSIRIS-REx will thoroughly characterize asteroid Bennu providing knowledge of the nature of near-Earth asteroids that is fundamental to understanding planet formation and the origin of life. The return to Earth of pristine samples with known geologic context will enable precise analyses that cannot be duplicated by spacecraft-based instruments, revolutionizing our understanding of the early Solar System. Bennu is both the most accessible carbonaceous asteroid and one of the most potentially Earth-hazardous asteroids known. Study of Bennu addresses multiple NASA objectives to understand the origin of the Solar System and the origin of life and will provide a greater understanding of both the hazards and resources in near-Earth space, serving as a precursor to future human missions to asteroids.This paper focuses on the technical aspects of the Sample Return Capsule (SRC) design and concept of operations, including trajectory design and reentry retrieval. Highlights of the mission are included below.The OSIRIS-REx spacecraft provides the essential functions for an asteroid characterization and sample return mission: attitude control propulsion power thermal control telecommunications command and data handling structural support to ensure successful rendezvous with Bennu characterization of Bennus properties delivery of the sampler to the surface, and return of the spacecraft to the vicinity of the Earth sample collection, performed by the Touch-and-Go Sample Acquisition Mechanism (TAGSAM), to acquire a regolith sample from the surface Earth re-entry and SRC recovery. Following sample collection, OSIRIS-REx drifts away from Bennu until the Asteroid Departure Maneuver is commanded on March 4, 2021, sending OSIRIS-REx on a ballistic return cruise to Earth. No additional large deterministic maneuvers are required to return the SRC to Earth. During the cruise, tracking and trajectory correction maneuvers (TCMs) are performed as necessary to precisely target the entry corridor. As OSIRIS-REx approaches Earth, the reentry plans are reviewed starting about a year before arrival, and preparations begin. The spacecraft is targeted away from the Earth until 7 days before entry. The final two trajectory correction maneuvers bring the spacecraft on target toward the Utah Test and Training Range (UTTR), with sufficient time for contingency resolution. The SRC releases 4 hours prior to atmospheric entry interface and, using the Stardust capsule heritage design, employs a traditional drogue and main parachute descent system for a soft touchdown.

reentry↗
NASA NTRS1997

Navigation and the Mars global surveyor mission

Navigation and trajectory planning aspects for the Mars global surveyor (MGS) mission, are discussed. In the first 205 days of the mission since launch on 7 November July 1996, two trajectory maneuvers were successfully completed. The third maneuver was cancelled due to the accuracy of the targeting of the previous maneuvers. The final trajectory correction maneuver will occur on arrival in order to adjust the arrival conditions for the Mars orbit insertion maneuver. The planned capture orbit is highly elliptical with a 45 hour orbital period, a 300 km periapsis altitude and a descending node. Two-way coherent Doppler and time delay measurements are acquired in the X-band by the Deep Space network and used to navigate the spacecraft. After capture, the orbit will be circularized by aerobraking and propulsive maneuvers. The three-phase aerobraking maneuver is described. During the 687 earth-day observation period, navigation will maintain a 2:00 pm descending node, sun-synchronous low altitude, short periodic orbit. Ground track coverage will remain almost uniform in order to maximize data acquisition.

Esposito, P.↗