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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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At least 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Imaging on ballistic missions to comet Encke

Imaging for navigation and science has been studied for a 1980 Encke flyby at 0.4 to 0.8 AU from the sun with spinning and three-axis spacecraft. Trajectory errors, maneuvers, encounter geometry, imaging performance, and data transmission were considered. Onboard comet sightings are needed for navigation. Recommended for a three-axis spacecraft are two vidicon cameras and a nucleus sensor for closed-loop pointing control. The cameras are essentially the Mariner 9 and Mariner 10 instruments; the nucleus sensor is an updated version of a sensor flown on Mariners 6 and 7. For a spinning spacecraft, a framing camera using a charge-coupled device and a spin-scan photometer are proposed. The framing camera would be a new design; it would not be despun, but the spin axis should point along the comet-centered velocity vector. The photometer is essentially that flown on Pioneer 10 and 11.-

Jaffe, L. D.↗
NASA NTRS2007

Space Technology 7 : Micropropulsion and Mass Distribution

The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The ST7 disturbance reduction system (DRS) will contain new micropropulsion technology to be flown as part of the European Space Agency's LISA (laser interferometer space antenna) Pathfinder project. After launch into a low Earth orbit in early 2010, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control position reference will be provided by the European LISA Technology Package, which will include two nearly free-floating test masses. The test mass position and attitude will be sensed and adjusted using electrostatic capacitance bridges. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom, six for each of the test masses and six for the spacecraft. In the absence of other disturbances, the test masses will slowly gravitate toward local concentrations of spacecraft mass. The test mass acceleration must be minimized to maintain the acceleration of the enclosing drag-free spacecraft within the control authority of the micropropulsion system. Therefore, test mass acceleration must be predicted by accurate measurement of mass distribution, then offset by the placement of specially shaped balance masses near each test mass. The - acceleration is characterized by calculating the gravitational effect of over ten million modeled points of a nearly 500-kg spacecraft. This paper provides an overview of the mission technology and the process of precision mass modeling of the DRS equipment.

ST-7↗
NASA NTRS1992

On the optimal two-impulse maneuver

An algorithm and a solution to the optimal two-impulse maneuver for the Ulysses spacecraft is presented. The option in which both maneuvers are designed using the critical plane strategy, which further reduces the propellant budget, is included. Results show savings of 12.4 and 16.7 m/s for full targeting and critical plane strategy, respectively. It is argued that this software will play an important role in the area of multimission maneuver design.

Bonneau, F. G.↗
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 NTRS1988

'Dogleg' maneuvers for Project Galileo

This paper describes the analysis and design of 'dog-leg' maneuvers for the spin-stabilized spacecraft of Project Galileo. These maneuvers involve a turn from one attitude to a firing attitude and then, following the motor burn, a turn to a new attitude. In general, the three attitudes (initial, firing and final) are noncoplanar. A computer program has been developed to find the firing attitude for the optimum conservation of propellant, while attaining a desired velocity change for the spacecraft. Results for several cases are presented.

Stanford, Richard H.↗
NASA NTRS2010

Implementation of an Autonomous Multi-Maneuver Targeting Sequence for Lunar Trans-Earth Injection

Using a fully analytic initial guess estimate as a first iterate, a targeting procedure that constructs a flyable burn maneuver sequence to transfer a spacecraft from any closed Moon orbit to a desired Earth entry state is developed and implemented. The algorithm is built to support the need for an anytime abort capability for Orion. Based on project requirements, the Orion spacecraft must be able to autonomously calculate the translational maneuver targets for an entire Lunar mission. Translational maneuver target sequences for the Orion spacecraft include Lunar Orbit Insertion (LOI), Trans-Earth Injection (TEI), and Trajectory Correction Maneuvers (TCMs). This onboard capability is generally assumed to be supplemental to redundant ground computation in nominal mission operations and considered as a viable alternative primarily in loss of communications contingencies. Of these maneuvers, the ability to accurately and consistently establish a flyable 3-burn TEI target sequence is especially critical. The TEI is the sole means by which the crew can successfully return from the Moon to a narrowly banded Earth Entry Interface (EI) state. This is made even more critical by the desire for global access on the lunar surface. Currently, the designed propellant load is based on fully optimized TEI solutions for the worst case geometries associated with the accepted range of epochs and landing sites. This presents two challenges for an autonomous algorithm: in addition to being feasible, the targets must include burn sequences that do not exceed the anticipated propellant load.

Whitley, Ryan J.↗
NASA NTRS2004

Dynamics and Controls of a Conceptual Jovian Moon Tour Spacecraft

The dynamics and control challenges presented by a conceptual Jovian Moon Tour spacecraft are summarized in this paper. Attitude and orbital dynamics interactions are present due to the designed low-thrust trajectory, and controls structure interactions are also present due to the non-collocated sensor-actuator pairs on board the flexible spacecraft. A finite-element based simulation model is described which is capable of handling the complex orbital and attitude dynamics arising during the low-thrust spiraling maneuvers of the spacecraft. A few numerical simulations demonstrate that some of the challenges hitherto identified can be faced via integrated dynamics and control analysis, and that reasonable assessments of the pointing performance can be made.

dynamics↗
NASA NTRS1988

Flexible Docking Tunnel

Brief report proposes use of flexible tunnel to provide "soft docking" between Space Shuttle Orbiter and space station during resupply visits. Soft docking reduces risks of structural damage to either spacecraft during docking maneuvers, provides adaptable docking interface, and avoids jarring sensitive material-processing experiments in space station. After docking, tunnel pressurized to enable exchange of people and cargo between two spacecraft.

Michaud, Roger↗
NASA NTRS1992

Fuel optimal maneuver: Experimental testbeds

The objective of this paper is to conduct experiments in fuel optimal propulsive maneuvers of flexible spacecraft in order to verify and extend recently developed theory and apply it to various classes of spacecraft. Topics covered in viewgraph format include: impulse damping control by experimental verification; impulse damping of the SCOLE reflector and mast; fuel optimal control of rotating flexible beam; fuel optimal aerobrake maneuvers.

Meyer, John L.↗
NASA NTRS1984

Identification and control of structures in space

The derivation of the equations of motion for the Spacecraft Control Laboratory Experiment (SCOLE) is reported and the equations of motion of a similar structure orbiting the earth are also derived. The structure is assumed to undergo large rigid-body maneuvers and small elastic deformations. A perturbation approach is proposed whereby the quantities defining the rigid-body maneuver are assumed to be relatively large, with the elastic deformations and deviations from the rigid-body maneuver being relatively small. The perturbation equations have the form of linear equations with time-dependent coefficients. An active control technique can then be formulated to permit maneuvering of the spacecraft and simultaneously suppressing the elastic vibration.

Meirovitch, L.↗
NASA NTRS1977

The dynamics and optimal control of spinning spacecraft and movable telescoping appendages, part B: Effect of gravity-gradient torques on the dynamics of a spinning spacecraft with telescoping appendages

The effects of gravity gradient torques during boom deployment maneuvers of a spinning spacecraft are examined. Configurations where the booms extended only along the hub principal axes and where one or two booms are offset from the principal axes were considered. For the special case of symmetric deployment (principal axes booms) the stability boundaries are determined, and a stability chart is used to study the system behavior. Possible cases of instability during this type of maneuver are identified. In the second configuration an expression for gravity torque about the hub center of mass was developed. The nonlinear equations of motion are solved numerically, and the substantial influence of the gravity torque during asymmetric deployment maneuvers is indicated.

Bainum, P. M.↗
NASA NTRS2019

Cloud Computing Methods for Near Rectilinear Halo Orbit Trajectory Design

Complicated mission design problems require innovative computational solutions. As spacecraft depart from a proposed Gateway in a Near Rectilinear Halo Orbit (NRHO), recontact analysis is required to avoid risk of collision and ensure safe operations. Escape dynamics from NRHOs are governed by multiple gravitational bodies, yielding a trajectory design space that is exhaustively large. This paper summarizes the recontact analysis for departure from the NRHO and describes how the Deep Space Trajectory Explorer (DSTE) trajectory design software incorporates high performance cloud computing to compute and visualize the orbit design space. Recent focus on exploration missions to cislunar space has kindled accelerated interest in multibody orbit solutions. Trajectory analysis in the presence of multiple gravity fields is complex, and innovative computational tools are needed to simplify complicated design spaces, to generate large quantities of data quickly, and to visualize the output for user accessibility. The Gateway mission is a prime example. The Gateway1 is proposed as a human outpost in deep space. The current baseline orbit for the Gateway is a Near Rectilinear Halo Orbit (NRHO) near the Moon.2 The NRHO exists in a regime that experiences the gravitational effects of the Earth and the Moon simultaneously, complicating orbit analysis. The mission design process benefits greatly from updated computational tools for multibody missions like the Gateway. As an example, consider the problem of assessing the risk of collision in an NRHO. As a staging location to missions to the lunar surface and beyond the Earth-Moon system, the Gateway will experience spacecraft and other objects regularly arriving and departing. Departing objects potentially include spent logistics modules, visiting crew vehicles, debris objects, wastewater particles, and cubesats. Each departure is governed by the dynamics of the Gateway orbit and the surrounding dynamical environment. Over time, any unmaintained object in such an orbit eventually departs due to the small instabilities associated with the NRHOs. A separation maneuver speeds the departure from the NRHO, but the effects of the maneuver on the spacecraft behavior depend on the location, magnitude, and direction of the burn. Escape dynamics from the NRHO with regard to these maneuver options open up an enormous potential trajectory design space where subtle changes in input can produce dramatically large changes in the results. Any departing object must avoid recontacting the Gateway as it leaves the lunar vicinity, and a recontact analysis thus involves a significant number of computations and extensive output data. To explore the dynamics of this extensive design space, the Deep Space Trajectory Explorer3 (DSTE) trajectory design software incorporates new High Performance Computing (HPC) services and novel interactive visualizations. This paper details the HPC and cloud infrastructure techniques that are implemented in the DSTE, applying the new capabilities to analysis of recontact risk with the Gateway in NRHO. NEAR RECTILINEAR HALO ORBITS The Gateway is planned to fly in a lunar NRHO as its baseline orbit. The NRHO families of orbits are subsets of the larger halo families, which originate from planar orbits near the L1 and L2 libration points; the Earth-Moon L2 halo family appears in Figure 1. Each halo orbit is perfectly periodic in the Circular Restricted 3-Body Problem (CR3BP) and becomes a quasi-periodic orbit in a higher fidelity ephemeris force model. The NRHOs are defined as those members of the halo family with bounded stability properties;2 they pass near the Moon at perilune and are nearly polar. Families exist with apolunes located both above the lunar north pole and above the lunar south pole; the Gateway is planned to reside in a southern L2 NRHO in a 9:2 resonance with the lunar synodic period. The 9:2 NRHO is characterized by a period of about 6.5 days, a perilune radius of about 3,500 km, and an apolune radius of about 71,000 km; it is strongly affected by the gravity of both the Earth and the Moon simultaneously. This NRHO offers extended communications with assets on the south pole of the Moon,4 as well as low-cost orbit maintenance and attitude control,5 favorable eclipse avoidance properties,6 and inexpensive transfers from Earth and to other destinations.5,7 The NRHO portion of the southern L2 halo family is highlighted in black in Figure 1, and the 9:2 NRHO appears in blue.

Phillips, Sean M.↗
NASA NTRS1989

NASA CSI suspension methods overview

New suspension techniques will be necesary for ground testing the flexible spacecraft anticipated in NASA's future space activity. The most complex spacecraft involve nonlinear maneuvering (i.e., large angle slewing) with articulating substructures such as remote manipulating systems. The NASA control-structure interaction (CSI) ground test method team has begun researching and developing methodology to suspend the future class of spacecraft. This overview describes the work completed thus far. The research objective and technical approach will be presented first. Second, will be a suspension device overview followed by an assessment of existing hardware. Two different mechanical zero-spring-rate mechanisms will be compared for optimal performance. Next, will be a description of how existing hardware can be evolved to meet more general suspension requirements. A comparison of suspending articulating structures overhead vs underneath will follow. After a few experimental results from the zero-spring-rate mechanism/air suspension cart will be concluding remarks and future work.

Woodard, Stanley E.↗
NASA NTRS2012

New Mission Old Spacecraft: EPOXI's Approach to the Comet Hartley-2

NASA's Deep Impact mission ended successfully in 2005 after an impact and close flyby of the comet 9P/Tempel-1. The Flyby spacecraft was placed in hibernation and was left to orbit the sun. In 2007, engineers at the Jet Propulsion Laboratory brought the spacecraft out of hibernation and successfully performed two additional missions. These missions were EPOCh, Extra-solar Planetary Observation and Characterization, a photometric investigation of transiting exo-planets, and DIXI, Deep Impact eXtended Investigation, which maneuvered the Flyby spacecraft towards a close encounter with the comet 103P/Hartley- 2 on 4 November 2010. The names of these two scientific investigations combine to form the overarching mission's name, EPOXI. The encounter with 103P/Hartley-2 was vastly different from the prime mission's encounter with 9P/Tempel-1. The geometry of encounter was nearly 180 ? different and 103P/Hartley-2 was approximately one-quarter the size of 9P/Tempel-1. Mission operations for the comet flyby were broken into three phases: a) Approach, b) Encounter, and c) Departure. This paper will focus on the approach phase of the comet encounter. It will discuss the strategies used to decrease both cost and risk while maximizing science return and some of the challenges experienced during operations.

science return↗
NASA NTRS1986

Nonlinear attitude control of flexible spacecraft under disturbance torque

A control law for large-angle single-axis rotational maneuvers of a spacecraft-beam-tip body (an antenna or a reflector) configuration is presented. It is assumed that an unknown but bounded disturbance torque is acting on the spacecraft. A model reference adaptive torque control law is derived for the slewing of the space vehicle. This controller includes a dynamic system in the feedback path and requires only attitude angle and rate of the space vehicle for feedback. For damping out the elastic motion excited by the slewing maneuver, a stabilizer is designed assuming that a torquer and a force actuator are available at the tip body. The stabilizer uses only the flexible modes for the synthesis of the control law. Simulation results are presented to show that fast, large-angle rotational maneuvers can be performed using the adaptive controller and the stabilizer in spite of the presence of continuously acting unknown torque on the spacecraft.

Singh, Sahjendra N.↗
NASA NTRS1991

Proximity operations considerations affecting spacecraft design

Proximity operations can be defined as the maneuvering of two or more spacecraft within 1 nautical mile range, with relative velocity less than 10 feet per second. The passive vehicle is nontranslating and should provide for maintenance of the desired approach attitude. It must accommodate the active (translating) vehicle induced structural loads and performance characteristics (mating hardware tolerances), and support sensor compatibility (transponder, visual targets, etc.). The active vehicle must provide adequate sensor systems (relative state information, field-of-view, redundancy), flight control hardware (thruster sizing, minimal cross-coupling, performance margins, redundancy) and software (reconfigurable, attitude/rate modes, translation and rotation fine control authority) characteristic, and adequate non-propulsive consumables such as power. Operational concerns must be considered. These include the following: (1) the desired approach trajectory and relative orientation; (2) the active vehicle thruster plume effects (forces, torques, contamination) on the passive vehicle; and (3) procedures for contingencies such as loss of communications, sensor or propulsion failures, and target vehicle loss of control.

Staas, Steven K.↗
NASA NTRS2015

Precision Closed-Loop Orbital Maneuvering System Design and Performance for the Magnetospheric Multiscale Formation

NASAs Magnetospheric Multiscale (MMS) mission successfully launched on March 13,2015 (UTC) consists of four identically instrumented spin-stabilized observatories that function as a constellation to study magnetic reconnection in space. The need to maintain sufficiently accurate spatial and temporal formation resolution of the observatories must be balanced against the logistical constraints of executing overly-frequent maneuvers on a small fleet of spacecraft. These two considerations make for an extremely challenging maneuver design problem. This paper focuses on the design elements of a 6-DOF spacecraft attitude control and maneuvering system capable of delivering the high-precision adjustments required by the constellation designers specifically, the design, implementation, and on-orbit performance of the closed-loop formation-class maneuvers that include initialization, maintenance, and re-sizing. The maneuvering control system flown on MMS utilizes a micro-gravity resolution accelerometer sampled at a high rate in order to achieve closed-loop velocity tracking of an inertial target with arc-minute directional and millimeter-per second magnitude accuracy. This paper summarizes the techniques used for correcting bias drift, sensor-head offsets, and centripetal aliasing in the acceleration measurements. It also discusses the on-board pre-maneuver calibration and compensation algorithms as well as the implementation of the post-maneuver attitude adjustments.

Maneuver and Attitude Dynamics↗
NASA NTRS2015

Precision Closed-Loop Orbital Maneuvering System Design and Performance for the Magnetospheric Multi-Scale Mission (MMS) Formation

NASA's Magnetospheric Multi-Scale (MMS) mission successfully launched on March 13, 2015 (UTC) consists of four identically instrumented spin-stabilized observatories that function as a constellation to study magnetic reconnection in space. The need to maintain sufficiently accurate spatial and temporal formation resolution of the observatories must be balanced against the logistical constraints of executing overly-frequent maneuvers on a small fleet of spacecraft. These two considerations make for an extremely challenging maneuver design problem. This paper focuses on the design elements of a 6-DOF spacecraft attitude control and maneuvering system capable of delivering the high-precision adjustments required by the constellation designers---specifically, the design, implementation, and on-orbit performance of the closed-loop formation-class maneuvers that include initialization, maintenance, and re-sizing. The maneuvering control system flown on MMS utilizes a micro-gravity resolution accelerometer sampled at a high rate in order to achieve closed-loop velocity tracking of an inertial target with arc-minute directional and millimeter-per-second magnitude accuracy. This paper summarizes the techniques used for correcting bias drift, sensor-head offsets, and centripetal aliasing in the acceleration measurements. It also discusses the on-board pre-maneuver calibration and compensation algorithms as well as the implementation of the post-maneuver attitude adjustments.

Determination and Control↗