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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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NASA NTRS2002

Orbit Determination Issues for Libration Point Orbits

Libration point mission designers require knowledge of orbital accuracy for a variety of analyses including station keeping control strategies, transfer trajectory design, and formation and constellation control. Past publications have detailed orbit determination (OD) results from individual notation point missions. This paper collects both published and unpublished results from four previous notation point missions (ISEE-3, SOHO, ACE and MAP) supported by Goddard Space Flight Center's Guidance, Navigation & Control Center. The results of those missions are presented along with OD issues specific to each mission. All past missions have been limited to ground based tracking through NASA ground sites using standard marine and Doppler measurement types. Advanced technology is enabling other OD options including onboard navigation using onboard attitude sensors and the use of the Very Long Baseline Interferometry (VLBI) measurement Delta Differenced One-Way Range (DDOR). Both options potentially enable missions to reduce coherent dedicated tracking passes while maintaining orbital accuracy. With the increased projected loading of the DSN, missions must find alternatives to the standard OD scenario.

Beckman, Mark↗

NASA NTRS1979

Concepts for generating optimum vertical flight profiles

Algorithms for generating optimum vertical profiles are derived and examined. These algorithms form the basis for the design of onboard flight management concepts. The variations in the optimum vertical profiles (resulting from these concepts) due to variations in wind, takeoff weight, and range-to-destination are presented. Further considerations for mechanizing two different onboard methods of computing near-optimum flight profiles are then outlined. Finally, the results are summarized, and recommendations are made for further work. Technical details of optimum trajectory design, steering requirements for following these trajectories, and off-line computer programs for testing the concepts are included.

Sorensen, J. A.↗

NASA NTRS2008

Descent Assisted Split Habitat Lunar Lander Concept

The Descent Assisted Split Habitat (DASH) lunar lander concept utilizes a disposable braking stage for descent and a minimally sized pressurized volume for crew transport to and from the lunar surface. The lander can also be configured to perform autonomous cargo missions. Although a braking-stage approach represents a significantly different operational concept compared with a traditional two-stage lander, the DASH lander offers many important benefits. These benefits include improved crew egress/ingress and large-cargo unloading; excellent surface visibility during landing; elimination of the need for deep-throttling descent engines; potentially reduced plume-surface interactions and lower vertical touchdown velocity; and reduced lander gross mass through efficient mass staging and volume segmentation. This paper documents the conceptual study on various aspects of the design, including development of sortie and outpost lander configurations and a mission concept of operations; the initial descent trajectory design; the initial spacecraft sizing estimates and subsystem design; and the identification of technology needs

Mazanek, Daniel D.↗

NASA NTRS2018

Overview of the Spacecraft Design for the Psyche Mission Concept

In January 2017, Psyche and a second mission concept were selected by NASA for flight as part of the 14th Discovery mission competition. Assigned for an initial launch date in 2023, the Psyche team was given direction shortly after selection to research the possibility for earlier opportunities. Ultimately, the team was able to identify a launch opportunity in 2022 with a reduced flight time to its destination. This was accomplished in large part to crosscutting trades centered on the electrical power subsystem. These trades were facilitated through the Psyche mission's planned use of Solar Electric Propulsion (SEP), which enables substantial flexibility with respect to trajectory design. In combination with low-thrust trajectory analysis tools, the team was able to robustly converge to solutions with a higher fidelity and accuracy of results. These trades also took advantage of the 1300 series product line produced by Space Systems Loral (SSL), which enabled power growth while maintaining strong system-level heritage through its modular design that has been utilized on a large number of geostationary (GEO) communications satellites. This paper presents an overview of the Psyche mission concept, and the unique architecture that enables the use of commercially developed electric propulsion and space power systems from Space Systems Loral to provide flexibility in mission design. This paper then discusses the trades that allowed the Psyche team to meet a 2022 launch date.

Prikl, Zachary↗

NASA NTRS1972

Gravity thrust Jupiter orbiter trajectories generated by encountering the Galilean satellites.

A trajectory design philosophy is introduced for Jupiter orbiter missions that is based on the gravity thrust concept developed for interplanetary trajectories. This is accomplished by utilizing the moving gravitational fields of the four Galilean satellites. A general theorem, applicable to any existing planetary satellite system, is developed to show how the effects of small satellite perturbations can be magnified by incorporating multiple revolutions between encounters.-

Minovitch, M. A.↗

NASA NTRS2009

ARTEMIS: The First Mission to the Lunar Libration Orbits

The ARTEMIS mission will be the first to navigate to and perform stationkeeping operations around the Earth-Moon L1 and L2 Lagrangian points. The NASA Goddard Space Flight Center (GSFC) has previous mission experience flying in the Sun-Earth L1 (SOHO, ACE, WIND, ISEE-3) and L2 regimes (WMAP) and have maintained these spacecraft in libration point orbits by performing regular orbit stationkeeping maneuvers. The ARTEMIS mission will build on these experiences, but stationkeeping in Earth-Moon libration orbits presents new challenges since the libration point orbit period is on the order of two weeks rather than six months. As a result, stationkeeping maneuvers to maintain the Lissajous orbit will need to be performed frequently, and the orbit determination solutions between maneuvers will need to be quite accurate. The ARTEMIS mission is a collaborative effort between NASA GSFC, the University of California at Berkeley (UCB), and the Jet Propulsion Laboratory (JPL). The ARTEMIS mission is part of the THEMIS extended mission. ARTEMIS comprises two of the five THEMIS spacecraft that will be maneuvered from near-Earth orbits into lunar libration orbits using a sequence of designed orbital maneuvers and Moon & Earth gravity assists. In July 2009, a series of orbit-raising maneuvers began the proper orbit phasing of the two spacecraft for the first lunar flybys. Over subsequent months, additional propulsive maneuvers and gravity assists will be performed to move each spacecraft though the Sun-Earth weak stability regions and eventually into Earth-Moon libration point orbits. We will present the overall orbit designs for the two ARTEMIS spacecraft and provide analysis results of the 3/4-body dynamics, and the sensitivities of the trajectory design to both · maneuver errors and orbit determination errors. We will present results from the. initial orbit-raising maneuvers.

Woodward, Mark↗

NASA NTRS2005

Analysis of Thrust Vectoring Capabilities for the Jupiter Icy Moons Orbiter

A strategy to mitigate the impact of the trajectory design of the Jupiter Icy Moons Orbiter (JIMO) on the attitude control design is described in this paper. This paper shows how the thrust vectoring control torques, i.e. the torques required to steer the vehicle, depend on various parameters (thrust magnitude, thrust pod articulation angles, and thrust moment arms). Rather than using the entire reaction control system (RCS) system to steer the spacecraft, we investigate the potential utilization of only thrust vectoring of the main ion engines for the required attitude control to follow the representative trajectory. This study has identified some segments of the representative trajectory where the required control torque may exceed the designed ion engine capability, and how the proposed mitigation strategy succeeds in reducing the attitude control torques to within the existing capability.

low thrust↗

NASA NTRS1983

Satellite tour design for the Galileo mission

The design of Galileo's tour of Jupiter's Galilean satellites is discussed. The Galileo mission is reviewed, and the gravity-assist trajectory design is described. Attention is given to mission constraints and science requirements, several tour design strategies, and a strawman satellite tour. The software used to design a satellite tour is examined.

Diehl, R. E.↗

NASA NTRS1979

Viking Navigation: Introduction

The design, control, and reconstruction of the flight paths of all four Viking vehicles is presented. The specification of requirements on the flight hardware is described along with the navigation strategies, procedures, operational software, and the inflight navigation. Features of the Viking navigation discussed include the precise determination of the spacecraft trajectories, prediction of the trajectories, design of the propulsive maneuvers required to effect the necessary trajectory changes, and calculation of the Lander descent guidance parameters.

Oneil, W. J.↗

NASA NTRS1990

Return to the moon - The Lunar Observer mission

The Lunar Observer mission seeks to understand the moon in its global context. Motivation for this project comes from scientific interest and the need for robotic precursor missions in the proposed Space Exploration Initiative. The multiple roles to be performed by this mission mean that the trajectory design process promises to be quite complex and unique. The choice of launch vehicle, launch opportunity and daily launch window are affected by the selection of the translunar transfer trajectory, as well as the arrival conditions desired at lunar orbit insertion. The length of the mission and the tight lunar orbit control requirements determine the design of the circumlunar portions of the Lunar Observer mission.

Cook, Richard A.↗

NASA NTRS1995

Several Ways to Leave for Luna

This paper is a comprehensive survey of ballistic trajectory designs leading to the delivery of small spacecraft to the surface of the Moon. All of the currently known types of ballistic transfer trajectories are examined (including some new ones):.

Luna Earth↗

NASA NTRSDate not supplied

Clustering Approach To Identifying Low Lunar Frozen Orbits In A High-Fidelity Model

Low lunar frozen orbits are of continued interest in the astrodynamics community for trajectory design and space domain awareness. This paper presents a data-driven approach to analyzing a wide variety of numerically-generated lunar trajectories in a 100 × 100 lunar gravity model with the point mass gravity of the Earth and Sun. First, clustering is used to extract a summary of these trajectories: within each cluster, trajectories possess a geometrically similar evolution of perilune but varying drift and lifetimes. Within some clusters, trajectories with a bounded perilune evolution are also identified to produce candidates for lunar frozen orbits of distinct geometries.

lunar frozen orbit gravity model data clustering↗

NASA NTRS2003

Application of local Lyapunov exponents to maneuver design and navigation in the three-body problem

Dynamical systems theory has recently been employed for several missions to design trajectories within the three-body problem. This research applied a stability technique, the calculation of local Lyapunov exponents, to such trajectories. Local Lyapunov exponents give an indication of the effects that perturbations or maneuvers will have on trajectories over a specified time.

local Lyapunov exponent design navigation three-bo↗

NASA NTRS2003

Flight Reconstruction of the Mars Pathfinder Disk-Gap-Band Parachute Drag Coefficient

On July 4, 1997, the Mars Pathfinder (MPF) mission successfully landed on Mars. The entry, descent, and landing (EDL) scenario employed the use of a Disk-Gap-Band parachute design to decelerate the Lander. Flight reconstruction of the entry using MPF flight accelerometer data revealed that the MPF parachute decelerated faster than predicted. In the summer of 2003, the Mars Exploration Rover (MER) mission will send two Landers to the surface of Mars arriving in January 2004. The MER mission utilizes a similar EDL scenario and parachute design as that employed by MPF. As a result, characterizing the degree of underperformance of the MPF parachute system is critical for the MER EDL trajectory design. This paper provides an overview of the methodology utilized to estimate the MPF parachute drag coefficient as experienced on Mars.

Desai, Prasun↗

NASA NTRS2008

Constraint Force Equation Methodology for Modeling Multi-Body Stage Separation Dynamics

This paper discusses a generalized approach to the multi-body separation problems in a launch vehicle staging environment based on constraint force methodology and its implementation into the Program to Optimize Simulated Trajectories II (POST2), a widely used trajectory design and optimization tool. This development facilitates the inclusion of stage separation analysis into POST2 for seamless end-to-end simulations of launch vehicle trajectories, thus simplifying the overall implementation and providing a range of modeling and optimization capabilities that are standard features in POST2. Analysis and results are presented for two test cases that validate the constraint force equation methodology in a stand-alone mode and its implementation in POST2.

Toniolo, Matthew D.↗

NASA NTRS2015

ARTEMIS Lunar Orbit Insertion and Science Orbit Design Through 2013

As of late-July 2011, the ARTEMIS mission is transferring two spacecraft from Lissajous orbits around Earth-Moon Lagrange Point #1 into highly-eccentric lunar science orbits. This paper presents the trajectory design for the transfer from Lissajous orbit to lunar orbit insertion, the period reduction maneuvers, and the science orbits through 2013. The design accommodates large perturbations from Earth's gravity and restrictive spacecraft capabilities to enable opportunities for a range of heliophysics and planetary science measurements. The process used to design the highly-eccentric ARTEMIS science orbits is outlined. The approach may inform the design of future planetary moon missions.

lunar orbiters↗

NASA NTRS2010

Development of a Tool to Recreate the Mars Science Laboratory Aerothermal Environment

The Mars Science Laboratory will enter the Martian atmosphere in 2012 with multiple char depth sensors and in-depth thermocouples in its heatshield. The aerothermal environment experienced by MSL may be computationally recreated using the data from the sensors and a material response program, such as the Fully Implicit Ablation and Thermal (FIAT) response program, through the matching of the char depth and thermocouple predictions of the material response program to the sensor data. A tool, CHanging Inputs from the Environment of FIAT (CHIEF), was developed to iteratively change different environmental conditions such that FIAT predictions match within certain criteria applied to an external data set. The computational environment is changed by iterating on the enthalpy, pressure, or heat transfer coefficient at certain times in the trajectory. CHIEF was initially compared against arc-jet test data from the development of the MSL heatshield and then against simulated sensor data derived from design trajectories for MSL. CHIEF was able to match char depth and in-depth thermocouple temperatures within the bounds placed upon it for these cases. Further refinement of CHIEF to compare multiple time points and assign convergence criteria may improve accuracy.

Beerman, A. F.↗

NASA NTRS2006

Exploring Europa with an RPS-Powered Spacecraft Results of the Europa Explorer Mission Concept Study

This viewgraph presentation focuses on the results of the more recent and detailed Europa Explorer (EE) study. Based on the Europa Geophysical Explorer (EGE) the EE Study was more detailed and reached a modified design point, it re-affirmed all the conclusions reached during the EGE Study. The presentation reviews some of the important considerations of the study, including the trajectory design with earth gravity assists, the radiation considerations, the desired instruments for studying Europa, the total mass available, a conceptual illustration of the spacecraft. The attitude, propulsion and thermal control issues are also addressed. The data communications issues are reviewed. The expectations from the mission are summarized in the conclusion. These include a 90 day operational period, that is likely to continue for over a year; that EE would produce 1000 more observations than the Galileo mission; that EE would carry over 200 kg of instrumentation (including shielding); that EE would return over 21 Gigabits of data per Earth day; there would be about 340kg of unused mass, which could be used for more instrumentation, or a lander; and that this would be designed with currently available technology.

orbiter↗