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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2020

Power and Propulsion Element (PPE) Spacecraft Reference Trajectory Document

This document captures example reference trajectories for the PPE including a reference delivery orbit and orbit maintenance, an example cislunar orbit transfer and end-of-mission (EOM) disposal trajectory. The flexibility of electric propulsion offers, by its low thrust nature, multiple different trajectory options to transfer from one orbit to another. The trajectories captured in this document are representative examples of a low thrust transfer from the NRHO and to multiple cislunar orbits. This document provides a consistent set of data from mission design to be used in the design of the vehicle capable of flying the trajectory described. The data in this document will be used to create conference papers. In order to do so, we are ending this document through for external release.

Melissa L Mcguire↗
NASA NTRS2011

Ares I-X Test Flight Reference Trajectory Development

Ares I-X was the first test flight of NASA's Constellation Program's Ares I crew launch vehicle. Ares I is a two stage to orbit launch vehicle that provides crew access to low Earth orbit for NASA's future manned exploration missions. The Ares I first stage consists of a Shuttle solid rocket motor (SRM) modified to include an additional propellant segment and a liquid propellant upper stage with an Apollo J2X engine modified to increase its thrust capability. The modified propulsion systems were not available for the first test flight, thus the test had to be conducted with an existing Shuttle 4 segment reusable solid rocket motor (RSRM) and an inert Upper Stage. The test flight's primary objective was to demonstrate controllability of an Ares I vehicle during first stage boost and the ability to perform a successful separation. In order to demonstrate controllability, the Ares I-X ascent control algorithms had to maintain stable flight throughout a flight environment equivalent to Ares I. The goal of the test flight reference trajectory development was to design a boost trajectory using the existing RSRM that results in a flight environment equivalent to Ares I. A trajectory similarity metric was defined as the integrated difference between the Ares I and Ares I-X Mach versus dynamic pressure relationships. Optimization analyses were performed that minimized the metric by adjusting the inert upper stage weight and the ascent steering profile. The sensitivity of the optimal upper stage weight and steering profile to launch month was also investigated. A response surface approach was used to verify the optimization results. The analyses successfully defined monthly ascent trajectories that matched the Ares I reference trajectory dynamic pressure versus Mach number relationship to within 10% through Mach 3.5. The upper stage weight required to achieve the match was found to be feasible and varied less than 5% throughout the year. The paper will discuss the flight test requirements, provide Ares I-X vehicle background, discuss the optimization analyses used to meet the requirements, present analysis results, and compare the reference trajectory to the reconstructed flight trajectory.

Starr, Brett R.↗
NASA NTRS2019

White Paper: Gateway Destination Orbit Model: A Continuous 15 Year NRHO Reference Trajectory

The Gateway program has selected a Near-Rectilinear Halo Orbit (NRHO) of the L2 southern family as its operational orbit. In order to facilitate mission analysis for Gateway and associated programs utilizing this orbit, a representative reference trajectory 15 years in duration has been developed. The purpose of this white paper is to review the properties of this NRHO reference trajectory.

lunar synodic resonant↗
NASA NTRS1977

Space shuttle engineering and operations support. ALT separation reference trajectories for tailcone on orbiter forward and aft CG configurations. Mission planning, mission analysis and software formulation

A preflight analysis of the ALT separation reference trajectories for the tailcone on, forward, and aft cg orbiter configurations is documented. The ALT separation reference trajectories encompass the time from physical separation of the orbiter from the carrier to orbiter attainment of the maximum ALT interface airspeed. The trajectories include post separation roll maneuvers by both vehicles and are generated using the final preflight data base. The trajectories so generated satisfy all known separation design criteria and violate no known constraints. The requirement for this analysis is given along with the specifications, assumptions, and analytical approach used to generate the separation trajectories. The results of the analytical approach are evaluated, and conclusions and recommendations are summarized.

Glenn, G. M.↗
NASA NTRS2014

Updating the Reference Trajectory for the Cassini Solstice Mission

The Cassini-Huygens deep-space probe has successfully completed a four-year prime tour and a two-year extended tour of the Saturnian system. Now in a second extended phase called the Solstice Mission, the Cassini spacecraft will continue to gather data as directed by the reference trajectory until 2017. This paper will describe the process of how a reference trajectory update is prepared and delivered to the project by the navigation team during Solstice Mission flight operations. This paper will also document the timeline of products released and utilized, as well as the study to include an Enceladus occultation observation that occurs in 2016.

navigation operations↗
NASA NTRS2006

The Jupiter Icy Moons Orbiter reference trajectory

The proposed NASA Jupiter Icy Moons Orbiter (JIMO) mission would have used a single spacecraft to orbit Callisto, Ganymede, and Europa in succession. The enormous Delta-Velocity required for this mission (nearly [25 km/s]) would necessitate the use of very high efficiency electric propulsion. The trajectory created for the proposed baseline JIMO mission may be the most complex trajectory ever designed. This paper describes the current reference trajectory in detail and describes the processes that were used to construct it.

optimization↗
NASA NTRS2017

Overview of the Mission Design Reference Trajectory for NASA's Asteroid Redirect Robotic Mission

The National Aeronautics and Space Administration's (NASA's) recently cancelled Asteroid Redirect Mission was proposed to rendezvous with and characterize a 100 m plus class near-Earth asteroid and provide the capability to capture and retrieve a boulder off of the surface of the asteroid and bring the asteroidal material back to cislunar space. Leveraging the best of NASA's science, technology, and human exploration efforts, this mission was originally conceived to support observation campaigns, advanced solar electric propulsion, and NASA's Space Launch System heavy-lift rocket and Orion crew vehicle. The asteroid characterization and capture portion of ARM was referred to as the Asteroid Redirect Robotic Mission (ARRM) and was focused on the robotic capture and then redirection of an asteroidal boulder mass from the reference target, asteroid 2008 EV5, into an orbit near the Moon, referred to as a Near Rectilinear Halo Orbit where astronauts would visit and study it. The purpose of this paper is to document the final reference trajectory of ARRM and the challenges and unique methods employed in the trajectory design of the mission.

low thrust↗
NASA NTRS1976

Orbiter/carrier separation for the ALT free flight no. 1 reference trajectories. Mission planning, mission analysis and software formulation

Details of the generation of the separation trajectories are discussed. The analysis culminated in definition of separation trajectories between physical separation and orbiter/carrier vortex clearance. Specifications, assumptions and analytical approach used to generate the separation trajectories are presented. Results of the analytical approach are evaluated. Conclusions and recommendations are summarized. Supporting references are listed.

Glenn, G. M.↗
NASA NTRS2010

Altair Descent and Ascent Reference Trajectory Design and Initial Dispersion Analyses

The Altair Lunar Lander is the linchpin in the Constellation Program (CxP) for human return to the Moon. Altair is delivered to low Earth orbit (LEO) by the Ares V heavy lift launch vehicle, and after subsequent docking with Orion in LEO, the Altair/Orion stack is delivered through translunar injection (TLI). The Altair/Orion stack separating from the Earth departure stage (EDS) shortly after TLI and continues the flight to the Moon as a single stack. Altair performs the lunar orbit insertion (LOI) maneuver, targeting a 100-km circular orbit. This orbit will be a polar orbit for missions landing near the lunar South Pole. After spending nearly 24 hours in low lunar orbit (LLO), the lander undocks from Orion and performs a series of small maneuvers to set up for descending to the lunar surface. This descent begins with a small deorbit insertion (DOI) maneuver, putting the lander on an orbit that has a perilune of 15.24 km (50,000 ft), the altitude where the actual powered descent initiation (PDI) commences. At liftoff from Earth, Altair has a mass of 45 metric tons (mt). However after LOI (without Orion attached), the lander mass is slightly less than 33 mt at PDI. The lander currently has a single descent module main engine, with TBD lb(sub f) thrust (TBD N), providing a thrust-to-weight ratio of approximately TBD Earth g's at PDI. LDAC-3 (Lander design and analysis cycle #3) is the most recently closed design sizing and mass properties iteration. Upgrades for loss of crew (LDAC-2) and loss of mission (LDAC-3) have been incorporated into the lander baseline design (and its Master Equipment List). Also, recently, Altair has been working requirements analyses (LRAC-1). All nominal data here are from the LDAC-3 analysis cycle. All dispersions results here are from LRAC-1 analyses.

Kos, Larry D.↗
NASA NTRS2017

Precomputing Process Noise Covariance for Onboard Sequential Filters

Process noise is often used in estimation filters to account for unmodeled and mismodeled accelerations in the dynamics. The process noise covariance acts to inflate the state covariance over propagation intervals, increasing the uncertainty in the state. In scenarios where the acceleration errors change significantly over time, the standard process noise covariance approach can fail to provide effective representation of the state and its uncertainty. Consider covariance analysis techniques provide a method to precompute a process noise covariance profile along a reference trajectory using known model parameter uncertainties. The process noise covariance profile allows significantly improved state estimation and uncertainty representation over the traditional formulation. As a result, estimation performance on par with the consider filter is achieved for trajectories near the reference trajectory without the additional computational cost of the consider filter. The new formulation also has the potential to significantly reduce the trial-and-error tuning currently required of navigation analysts. A linear estimation problem as described in several previous consider covariance analysis studies is used to demonstrate the effectiveness of the precomputed process noise covariance, as well as a nonlinear descent scenario at the asteroid Bennu with optical navigation.

onboard↗
NASA NTRS1999

Martian Aerocapture Terminal Point Guidance: A Reference Path Optimization Study

An effective method of terminal point guidance is to employ influence coefficients, which are solved from a set of differential equations adjoint to the linearized perturbations of the equations of motion about a reference trajectory. Hence, to optimize this type of guidance, one must first optimize the reference trajectory that the guidance is based upon. This study concentrates on various methods to optimize a reference trajectory for a Martian aerocapture maneuver, including a parametric analysis and first order gradient method. Resulting reference trajectories were tested in separate 2000 6-DOF Monte Carlo runs, using the Atmospheric Guidance Algorithm Testbed for the Mars Surveyor Program 2001 (MSP '01) Orbiter. These results were compared to an August 1998 study using the same terminal point control guidance algorithm and simulation testbed. Satisfactory improvements over the 1998 study are amply demonstrated.

Ro, Theodore U.↗
DOE OSTI2020

Leveraging Multiple Connected Traffic Light Signals in an Energy-Efficient Speed Planner

Connecting automated vehicles to traffic lights can lead to significant energy savings by enabling them to pass through intersections in an energy-efficient way without unnecessary stops. A cellular-based communication system connecting multiple traffic lights can help realize the full potential of energy-efficient driving at intersections. Thus, we propose a hierarchical speed planner that can leverage information from multiple connected traffic lights. The proposed speed planner consists of two modules: a green window selector and a reference trajectory generator. The green window selector, based on Dijkstra’s algorithm, finds a series of "green windows" for connected traffic lights that builds an energy-optimal path for vehicles to follow. The reference trajectory generator finds optimal entering times, based on the selected green window at each intersection, and then computes reference trajectories. Deriving and using analytical optimal entering speeds as a function of entering times allows us to guarantee the computational simplicity suitable for real-time implementation. We also demonstrate how to balance energy and traffic flow perspectives in the reference trajectory generator. Lastly, a high-fidelity simulation framework is used to evaluate the proposed speed planner and quantify the extent to which it can save energy in various real-world urban route scenarios.

32 ENERGY CONSERVATION, CONSUMPTION, AND UTILIZATI↗
NASA NTRS2010

Preliminary Design of the Guidance, Navigation, and Control System of the Altair Lunar Lander

Guidance, Navigation, and Control (GN&C) is the measurement and control of spacecraft position, velocity, and attitude in support of mission objectives. This paper provides an overview of a preliminary design of the GN&C system of the Lunar Lander Altair. Key functions performed by the GN&C system in various mission phases will first be described. A set of placeholder GN&C sensors that is needed to support these functions is next described. To meet Crew safety requirements, there must be high degrees of redundancy in the selected sensor configuration. Two sets of thrusters, one on the Ascent Module (AM) and the other on the Descent Module (DM), will be used by the GN&C system. The DM thrusters will be used, among other purposes, to perform course correction burns during the Trans-lunar Coast. The AM thrusters will be used, among other purposes, to perform precise angular and translational controls of the ascent module in order to dock the ascent module with Orion. Navigation is the process of measurement and control of the spacecraft's "state" (both the position and velocity vectors of the spacecraft). Tracking data from the Earth-Based Ground System (tracking antennas) as well as data from onboard optical sensors will be used to estimate the vehicle state. A driving navigation requirement is to land Altair on the Moon with a landing accuracy that is better than 1 km (radial 95%). Preliminary performance of the Altair GN&C design, relative to this and other navigation requirements, will be given. Guidance is the onboard process that uses the estimated state vector, crew inputs, and pre-computed reference trajectories to guide both the rotational and the translational motions of the spacecraft during powered flight phases. Design objectives of reference trajectories for various mission phases vary. For example, the reference trajectory for the descent "approach" phase (the last 3-4 minutes before touchdown) will sacrifice fuel utilization efficiency in order to provide landing site visibility for both the crew and the terrain hazard detection sensor system. One output of Guidance is the steering angle commands sent to the 2 degree-of-freedom (dof) gimbal actuation system of the descent engine. The engine gimbal actuation system is controlled by a Thrust Vector Control algorithm that is designed taking into account the large quantities of sloshing liquids in tanks mounted on Altair. In this early design phase of Altair, the GN&C system is described only briefly in this paper and the emphasis is on the GN&C architecture (that is still evolving). Multiple companion papers will provide details that are related to navigation, optical navigation, guidance, fuel sloshing, rendezvous and docking, machine-pilot interactions, and others. The similarities and differences of GN&C designs for Lunar and Mars landers are briefly compared.

Lee, Allan Y.↗
NASA NTRS2001

A Non-Linear Approach to Spacecraft Trajectory Control in the Vicinity of a Libration Point

An expanding interest in mission design strategies that exploit libration point regions demands the continued development of enhanced, efficient, control algorithms for station-keeping and formation maintenance. This paper discusses the development of a non-linear, station-keeping, control algorithm for trajectories in the vicinity of a libration point. The control law guarantees exponential convergence, based on a Lyaponov analysis. Controller performance is evaluated using FreeFlyer(R) and MATLAB(R) for a spacecraft stationed near the L2 libration point in the Earth-Moon system, tracking a pre-defined reference trajectory. Evaluation metrics are fuel usage and tracking accuracy. Simulation results are compared with a linear-based controller for a spacecraft tracking the same reference trajectory. Although the analysis is framed in the context of station-keeping, the control algorithm is equally applicable to a formation flying problem with an appropriate definition of the reference trajectory.

Luquette, Richard J.↗