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Results for “interplanetary cruise”

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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

A Free-Return Earth-Moon Cycler Orbit for an Interplanetary Cruise Ship

A periodic circumlunar orbit is presented that can be used by an interplanetary cruise ship for regular travel between Earth and the Moon. This Earth-Moon cycler orbit was revealed by introducing solar gravity and modest phasing maneuvers (average of 39 m/s per month) which yields close-Earth encounters every 7 or 10 days. Lunar encounters occur every 26 days and offer the chance for a smaller craft to depart the cycler and enter lunar orbit, or head for a Lagrange point (e.g., EM-L2 halo orbit), distant retrograde orbit (DRO), or interplanetary destination such as a near-Earth object (NEO) or Mars. Additionally, return-to-Earth abort options are available from many points along the cycling trajectory.

Interplanetary cruise ship↗
NASA NTRS2007

Mars Reconnaissance Orbiter Interplanetary Cruise Navigation

Carrying six science instruments and three engineering payloads, the Mars Reconnaissance Orbiter (MRO) is the first mission in a low Mars orbit to characterize the surface, subsurface, and atmospheric properties with unprecedented detail. After a seven-month interplanetary cruise, MRO arrived at Mars executing a 1.0 km/s Mars Orbit Insertion (MOI) maneuver. MRO achieved a 430 km periapsis altitude with the final orbit solution indicating that only 10 km was attributable to navigation prediction error. With the last interplanetary maneuver performed four months before MOI, this was a significant accomplishment. This paper describes the navigation analyses and results during the 210-day interplanetary cruise. As of August 2007 MRO has returned more than 18 Terabits of scientific data in support of the objectives set by the Mars Exploration Program (MEP). The robust and exceptional interplanetary navigation performance paved the way for a successful MRO mission.

You, Tung-Han↗
NASA NTRS1984

The performance of differential VLBI delay during interplanetary cruise

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.

Moultrie, B.↗
NASA NTRS1984

The Performance of Differential VLBI Delay During Interplanetary Cruise

Project Voyager radio metric data are used to evaluate the orbit determination utilities 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 combination 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 was reduced by over 95% with a concurrent 90% reduction in the DSN time allocated to radio metric data acquisition.

Moultrie, B.↗
NASA NTRS1972

Mariner 9 propulsion subsystem performance during interplanetary cruise and Mars orbit insertion

On 14 November 1971 the Mariner 9 1334-N-(300-lbf)-thrust rocket engine was fired for just over 15 min to place the first man-made satellite into orbit about Mars. Propulsion subsystem data gathered during the 5-month interplanetary cruise and orbit insertion are of significance to future missions of this type. Specific results related to performance predictability, zero g heat transfer, and nitrogen permeation, diffusion, and solubility values are presented.

Cork, M. J.↗
NASA NTRS1992

Orbit-determination performance of Doppler data for interplanetary cruise trajectories. Part 1: Error analysis methodology

An error covariance analysis methodology is used to investigate different weighting schemes for two-way (coherent) Doppler data in the presence of transmission-media and observing-platform calibration errors. The analysis focuses on orbit-determination performance in the interplanetary cruise phase of deep-space missions. Analytical models for the Doppler observable and for transmission-media and observing-platform calibration errors are presented, drawn primarily from previous work. Previously published analytical models were improved upon by the following: (1) considering the effects of errors in the calibration of radio signal propagation through the troposphere and ionosphere as well as station-location errors; (2) modelling the spacecraft state transition matrix using a more accurate piecewise-linear approximation to represent the evolution of the spacecraft trajectory; and (3) incorporating Doppler data weighting functions that are functions of elevation angle, which reduce the sensitivity of the estimated spacecraft trajectory to troposphere and ionosphere calibration errors. The analysis is motivated by the need to develop suitable weighting functions for two-way Doppler data acquired at 8.4 GHz (X-band) and 32 GHz (Ka-band). This weighting is likely to be different from that in the weighting functions currently in use; the current functions were constructed originally for use with 2.3 GHz (S-band) Doppler data, which are affected much more strongly by the ionosphere than are the higher frequency data.

Ulvestad, J. S.↗
NASA NTRS1991

Galileo ultraviolet spectrometer experiment - Initial Venus and interplanetary cruise results

The Galileo Extreme Ultraviolet Spectrometer obtained a spectrum of Venus atmospheric emissions in the 55.0- to 125.0-nm wavelength region. Emissions of helium (58.4 nm), ionized atomic oxygen (83.4 nm), and atomic hydrogen (121.6 nm), as well as a blended spectral feature of atomic hydrogen (Lyman-beta) and atomic oxygen (102.5 nm), were observed at 3.5-nm resolution. During the Galileo spacecraft cruise from Venus to earth, Lyman-alpha emission from solar system atomic hydrogen (121.6 nm) was measured. The dominant source of the Lyman-alpha emission is atomic hydrogen from the interstellar medium. A model of Galileo observations at solar maximum indicates a decrease in the solar Lyman-alpha flux near the solar poles. A strong day-to-day variation also occurs with the 27-day periodicity of the rotation of the sun.

Hord, C. W.↗
NASA NTRS1992

Orbit-determination performance of Doppler data for interplanetary cruise trajectories. Part 2: 8.4-GHz performance and data-weighting strategies

A consider error covariance analysis was performed in order to investigate the orbit-determination performance attainable using two-way (coherent) 8.4-GHz (X-band) Doppler data for two segments of the planned Mars Observer trajectory. The analysis includes the effects of the current level of calibration errors in tropospheric delay, ionospheric delay, and station locations, with particular emphasis placed on assessing the performance of several candidate elevation-dependent data-weighting functions. One weighting function was found that yields good performance for a variety of tracking geometries. This weighting function is simple and robust; it reduces the danger of error that might exist if an analyst had to select one of several different weighting functions that are highly sensitive to the exact choice of parameters and to the tracking geometry. Orbit-determination accuracy improvements that may be obtained through the use of calibration data derived from Global Positioning System (GPS) satellites also were investigated, and can be as much as a factor of three in some components of the spacecraft state vector. Assuming that both station-location errors and troposphere calibration errors are reduced simultaneously, the recommended data-weighting function need not be changed when GPS calibrations are incorporated in the orbit-determination process.

Ulvestad, J. S.↗
NASA NTRS1992

Precision X-band radio Doppler and ranging navigation: Mars Observer interplanetary cruise scenario

This article describes an error covariance analysis based on a Mars Observer mission scenario; the study was performed to establish the navigation performance that can potentially be achieved in a demonstration of precision two-way X-band (8.4-GHz) Doppler and ranging with the Mars Observer spacecraft planned for next year, and to evaluate the sensitivity of the predicted performance to variations in ground system error modeling assumptions. Orbit determination error statistics computed for a 182-day Doppler and ranging data arc predicted Mars approach orbit determination accuracies of about 0.45 micro-rad in an angular sense, using a conservative ground system error model as a baseline. When less-conservative error model assumptions were employed, it was found that orbit determination accuracies of 0.19 to 0.30 micro-rad could be obtained; the level of accuracy of the assumed Mars ephemeris is about 0.11 micro-rad. In comparison, Doppler-only performance with the baseline error model was predicted to be about 1.30 to 1.51 micro-rad, although it was found that when improved station location accuracies and Global Positioning System-based tropospheric calibration accuracies were assumed, accuracies of 0.44 to 0.52 micro-rad were predicted. In the Doppler plus ranging cases, the results were relatively insensitive to variations in ranging system and station delay calibration uncertainties of a few meters and tropospheric zenith delay calibration uncertainties of a few centimeters.

Estefan, J. A.↗
NASA NTRS2001

Mars Global Surveyor Mars Orbiter Camera: Interplanetary Cruise through Primary Mission

More than three years of high-resolution (1.5-20 m/pixel) photographic observations of the surface of Mars have dramatically changed our view of that planet. Among the most important observations and interpretations derived therefrom are that much of Mars, at least to depths of several kilometers, is layered; that substantial portions of the planet have experienced burial and subsequent exhumation; that layered and massive units, many kilometers thick, appear to reflect an ancient period of large-scale erosion and deposition within what are now the ancient heavily cratered regions of Mars; and that processes previously unsuspected, including gully-forming fluid action and burial and exhumation of large tracts of land, have operated within near-contemporary times. These and many other attributes of the planet argue for a complex geology and complicated history.

Malin, Michael C.↗
NASA NTRS2001

Mars Global Surveyor Mars Orbiter Camera: Interplanetary Cruise Through Primary Mission

More than three years of high-resolution (1.5-20 m/pixel) photographic observations of the surface of Mars have dramatically changed our view of that planet. Among the most important observations and interpretations derived are that much of Mars, at least to depths of several kilometers, is layered; that substantial portions of the planet have experienced burial and subsequent exhumation; that layered and massive units, many kilometers thick, appear to reflect an ancient period of large-scale erosion and deposition within what are now the ancient heavily cratered regions of Mars; and that processes previously unsuspected, including gully-forming fluid action and burial and exhumation of large tracts of land, have operated within near-contemporary times. These and many other attributes of the planet argue for a complex geology and complicated history.

Malin, Michael C.↗
NASA NTRSDate not supplied

Aerocapture Trajectory Design for Uranus Orbiter

Introduction: The recently released National Academies Planetary Science and Astrobiology Decadal Survey 2023-2032 [1] identified the Ice Giants as the top priority science destination. While the survey acknowledged the potential for either a Uranus Orbiter and Probe (UOP) mission or a Neptune-Triton Odyssey mission, it ultimately identified the former as the highest priority new flagship mission. UOP missions calls for a launch window of opportunity between 2031-2038 with 12-15 year interplanetary cruise time along with a fully-propulsive Uranus Orbit Insertion burn on the order of a few km/s. However, a mission to Uranus with the same science payload could utilize aerocapture for orbit insertion to achieve both a significant reduction in the interplanetary cruise time and reduction in propulsive burn costs. Why Aerocapture: Aerocapture is a promising propellant and time-saving orbital insertion technique for planetary destinations with an atmosphere. Although not flight-proven, previous aerocapture systems studies in the literature have demonstrated both the validity and robustness of the technique at various planetary destinations. With respect to the Ice Giant planets, Neptune has seen more of the analysis in the literature. For science missions at Neptune, aerocapture can enable 1.4 times more delivered mass to orbit than an all-propulsive mission for the same launch vehicle while reducing interplanetary cruise times by more than 3 years [2] Additionally with modern guidance and control, Neptune aerocapture with blunt-body aeroshells is realizable [3][4]. There are limited papers in the literature investigating Uranus aerocapture with those available providing a preliminary feasibility assessment [5]. Consequently, the two-year funded NASA Space Technology Mission Directorate (STMD)-funded project, titled Aerocapture System as an Enabling Technology for Ice Giants Missions, aims to mature the analysis and technology state of Uranus aerocapture. Trajectory Design: This paper presents the current state of the trajectory design in support of the new aerocapture project. The project design philosophy is inspired from recent Neptune aerocapture studies, which employed modern guidance and control, in the sense that blunt-body aeroshells are analyzed. An assessment of the theoretical flight path angle corridor width is conducted for a range of ballistic coefficients and lift-to-drag ratios for both Space Launch System and Falcon Heavy Launch Vehicle interplanetary trajectory solutions. The results from the corridor width assessment provide an assessment of the aerocapture design-space and qualitative metrics on trajectory design considerations. The Program to Optimize Simulated Trajectories II (POST2) is utilized to run Monte Carlo simulations of Uranus aerocapture three-degree-of-freedom bank angle modulated trajectories using a closed-loop numerical-predictor corrector guidance algorithm. UranusGRAM 2021 is utilized as the atmospheric model [6]. A Uranus-developed aerodatabase, originally derived from Mars Science Laboratory (MSL), is utilized to provide vehicle aerodynamics over a wide range of hypersonic flow regimes. A MSL-derived 70 deg 4.5m diameter sphere-cone aeroshell that houses the UOP payload mass is assumed. Robustness testing and performance analysis is conducted, including the assessment of entry state errors, atmosphere density variations, and aerodynamic dispersions. Post-aerocapture Delta-V and aerothermal statistics are formulated into propellant mass and TPS requirements. The results presented in the paper will demonstrate the trajectory viability of Uranus aerocapture. Preliminary Results: Preliminary trajectory design results indicates successful Uranus aerocapture with a blunt-body aeroshell housing the same payload mass as the UOP mission from an Earth-to-Uranus interplanetary trajectory arriving in less than 7 years. From this interplanetary trajectory, aerocapture provides an orbit insertion Delta-V capability of 6.9 km/s requiring less than 300 m/s for post-aerocapture correction burns (15% of wet mass allocated to propellant). To put this into perspective, the UOP study utilized an Earth-to-Uranus trajectory that arrives in 13 years and requires more than 1000 m/s for fully-propulsive orbit insertion (40% of wet mass allocated to propellant). Achieving the same 6.9 km/s Delta-V capability fully-propulsively is mass prohibitive (97% of wet mass allocated to propellant). Nevertheless, aerocapture has the potential to reduce interplanetary transit times to Uranus by half while delivering the same payload mass to orbit in a reduced propellant mass footprint.

Rohan Deshmukh↗
NASA NTRS1983

Deep Space Navigation with Noncoherent Tracking Data

Navigation capabilities of noncoherent tracking data are evaluated for interplanetary cruise phase and planetary (Venus) flyby orbit determination. Results of a formal covariance analysis are presented which show that a combination of one-way Doppler and delta DOR yields orbit accuracies comparable to conventional two-way Doppler tracking. For the interplanetary cruise phase, a tracking cycle consisting of a 3-hour Doppler pass and delta DOR (differential one-way range) from two baselines (one observation per overlap) acquired 3 times a month results in 100-km orbit determination accuracy. For reconstruction of a Venus flyby orbit, 10 days tracking at encounter consisting of continuous one-way Doppler and delta DOR sampled at one observation per overlap is sufficient to satisfy the accuracy requirements.

Ellis, J.↗
NASA NTRS1986

ICE navigation support

The International Cometary Explorer (ICE) encounter with Comet Giacobini-Zinner took place 7 years after the spacecraft's original launch on 12 August 1978 as the International Sun Earth Explorer 3 (ISEE-3), part of a three-spacecraft project to study the interaction between the solar wind and the Earth's magnetosphere. Transfer to an interplanetary trajectory was performed via a 119-km-altitute, gravity-assist, lunar swingby on December 1983. Navigation support during interplanetary cruise and comet encounter was provided by orbit determination utilizing radio metric data from the DSN 64-meter antennas in Goldstone, California and Madrid, Spain. Orbit solutions yielding predictions of 50-km geocentric delivery accuracy in the target aim plane were achieved during interplanetary cruise and at comet encounter using 6-to-12-week data arcs between periodic attitude-change maneuvers. One-sigma two-way range and range rate residuals were consistently 40 meters and 0.2 mm/s or better, respectively. Non-gravitational forces affected the comet's motion during late August and early September 1985 and caused a 2300-km shift in the orbit of the comet relative to the spacecraft. This necessitated a final ICE orbit trim maneuver 3 days prior to encounter. Near-real-time assessment of two-way 2-GHz (S-band) Doppler pseudo-residuals during the June and July 1985 trajectory change maneuvers aided in calibration of the spacecraft's thrusters in preparation for this final critical maneuver. Post-flight analysis indicates tail centerline passage was achieved within 10 seconds of the predicted time and geocentric position uncertainty at encounter was less than 40 km.

Efron, L.↗
NASA NTRS2018

A Novel Multi-Spacecraft Interplanetary Global Trajectory Optimization Transcription

As the frontier of space exploration continues to advance, so does the design complexity of future interplanetary missions. One avenue of this increasing complexity includes a class of designs known as "Distributed Spacecraft Missions"; missions where multiple spacecraft coordinate to perform shared objectives. Current approaches for the global trajectory optimization of these Multi-Vehicle Missions (MVMs) are prone to shortcomings including laborious iterative design, considerable human-in-the-loop effort, treatment of the multi-vehicle problem as multiple separate trajectory optimization subproblems (resulting in suboptimal solutions where the whole is less than the sum of its parts), and poor handling of coordination objectives and constraints. There are only a handful of software platforms in existence capable of fully-automated, rapid, interplanetary mission and systems global optimization including the Parallel Global Multiobjective Optimizer (PaGMO), the Gravity Assisted Low-thrust Local Optimization Program (GALLOP), and the Evolutionary Mission Trajectory Generator (EMTG). However, none of these tools is capable of performing such tasks for MVM designs. The work outlined in this paper lays the groundwork for a technique to begin addressing these shortcomings. We present a fully-automated technique which frames interplanetary MVMs as Multi-Objective, Multi-Agent Hybrid Optimal Control Problems (MOMA HOCP). First, the basic functionality of this technique is validated on the single-vehicle problem of reproducing the Cassini interplanetary cruise.

Interplanetary↗