S-Band laboratory test equipment Final report
Subsystem component design and development, and system performance of ammonia fueled resistojet thrustor for spacecraft inversion maneuvers
RESISTOJET ENGINE↗
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Handling qualities for pilot control of Apollo lunar-landing spacecraft.
Simulated lunar landing maneuver of Apollo spacecraft, determining pilot control problems and handling qualities
ASTRONAUT PERFORMANCE↗
Dynamics of spacecraft control laboratory experiment (SCOLE) slew maneuvers
This is the first of two reports on the dynamics and control of slewing maneuvers of the NASA Spacecraft Control Laboratory Experiment (SCOLE). In this report, the dynamics of slewing maneuvers of SCOLE are developed in terms of an arbitrary maneuver about any given axis. The set of dynamical equations incorporate rigid-body slew maneuver and three-dimensional vibrations of the complete assembly comprising the rigid shuttle, the flexible beam, and the reflector with an offset mass. The analysis also includes kinematic nonlinearities of the entire assembly during the maneuver and the dynamics of the interaction between the rigid shuttle and the flexible appendage. The final set of dynamical equations obtained for slewing maneuvers is highly nonlinear and coupled in terms of the flexible modes and the rigid-body modes. The equations are further simplified and evaluated numerically to include the first ten flexible modes and the SCOLE data to yield a model for designing control systems to perform slew maneuvers.
Kakad, Y. P.↗
Radio occultation studies of the Venus atmosphere with the Magellan spacecraft. 1: Experimental description and performance
While primarily designed for radar studies of the Venus surface, the high effective isotropic radiated power (EIRP) from the Magellan spacecraft makes it an ideal transmitter for use in radio occultation measurements of the refractivity and absorptivity of the Venus atmosphere. Such experiments have been conducted involving transmissions at 2.3 GHz and 8.4 GHz (13 cm and 3.6 cm, respectively), during spacecraft ingress. Since the stability of the spacecraft transmitter is critical for accurately determining the Doppler shift and amplitude attenuation created as the ray penetrates the atmosphere, the spacecraft transmitter was locked to a 2.1 GHz uplink from a 70-meter DSN station which also received the signals. Because of the high directivity of the spacecraft antenna, and the significant ray bending in the deep Venus atmosphere, a spacecraft tracking maneuver was designed to keep the spacecraft antenna pointed in the direction of the refracted ray path back to Earth. This tracking maneuver, plus the high EIRP of the Magellan transmitter has yielded 3.6 cm refractivity and absorptivity profiles down to the 35 km altitude and 13 cm profiles down to the altitude of critical refraction (approximately 33 km). The statistical uncertainties in the derived profiles are significantly lower than those previously obtained, resulting in extremely accurate profiles of H2SO4 (g) abundance as discussed in an accompanying paper.
Steffes, Paul G.↗
Nonlinear slew maneuver dynamics of large flexible spacecrafts
In this paper, the dynamics of three-dimensional, large-angle arbitrary slew maneuvers of a large flexible spacecraft are developed. The dynamical equations obtained allow maneuver specifications about any axis and are highly nonlinear. They also include coupling between the rigid orbiter and the flexible appendage and correction for motion stiffness. A decentralized control scheme is utilized for performing the maneuver of the rigidized body and for vibration suppression of the flexible appendage. The method developed in this paper is further applied to NASA Spacecraft Control Laboratory Experiment (SCOLE) test facility.
Kakad, Y. P.↗
Wild2 approach maneuver strategy used for Stardust spacecraft (extended abstract)
Stardust, NASA's first dedicated sample return mission to a comet, successfully flew through the comet dust around Wold2 on January 2, 2004. The spacecraft flew within 236 km of the comet, meeting the mission requirement of 250+/- 50 km on flyby distance. Stardust collected dust particles and took several images of the comet while flying close to Wild2. The spacecraft will return to earth with the comet samples on January 15, 2006. To accomplish the above objective, a large Deep Space Maneuver (DSM#) was implemented during June 17 and 18, 2003 and a series of Trajectory Correction Maneuvers (TCMs) were also implemented during the 30 days prior to encounter. Both maneuver design and executions were influenced by number of factors including the small body ephemeris uncertainty, predictability of small forces arising from 3-axis attitude limit cycling and spacecraft slews. Maneuver design processes, including contingency plans, and maneuver performance characteristics, are discussed in this paper.
dust↗
Spacecraft attitude control during thrusting maneuvers - Space vehicle design criteria /Guidance and control/
Systems design criteria for spacecraft attitude control during thrusting maneuvers
Source record↗
Multibody Simulation of REMORA CubeSat Docking to and Pushing a Spent Rocket Booster
This paper details the multibody simulation of three phases of critical importance to the feasibility of the REMORA CubeSat space debris mitigation concept: the final approach of the CubeSat to a spent rocket booster; the grappling of the spent rocket booster using a robotic arm; and the pushing of the spent rocket booster to divert its course from another on-orbit asset. The extension of a robotic mobility and manipulation modeling toolkit (M3TK) from multibody dynamics simulation of manipulators and ground vehicles to simulation of orbital robotics is outlined. This includes the identification of the appropriate parameters required to concisely and generically describe thruster loads, thruster mixing, spacecraft control, and spacecraft navigation for the purpose of on-orbit robotics simulation. A high-level spacecraft navigator commands maneuvers to target spacecraft states. A PID spacecraft controller takes the target states and calculates desired forces and torques. A thruster mixer solves a quadratic program to determine the optimal thruster firing times for the propulsion system. Pulse width modulated actuation of eight canted cold gas thrusters is used in the simulated approach to a rocket nozzle from a distance of 200 m. A five degree of freedom robotic arm is controlled to position a pair of pincers to grasp the rocket nozzle. Contact dynamics are used to accurately simulate the grasping of the rocket nozzle by the pincers. A similar, but separate simulation is performed to assess the ability of the REMORA CubeSat to push the large spent rocket booster. This diversion maneuver makes use of an additional, larger thruster, and pushes the rocket booster in excess of 400 m. Appropriate motor control gains on the robotic arm are found to be higher during the pushing phase than those which are appropriate during free motion; this increase promotes rigidity of the arm and allows it to properly direct the pushing force. Challenges encountered in time step selection for numerical stability of the simulation are also discussed.
Mukherjee, Rudranarayan↗
Attitude stabilization of a rigid spacecraft using two momentum wheel actuators
It is well known that three momentum wheel actuators can be used to control the attitude of a rigid spacecraft and that arbitrary reorientation maneuvers of the spacecraft can be accomplished using smooth feedback. If failure of one of the momentum wheel actuators occurs, it is demonstrated that two momentum wheel actuators can be used to control the attitude of a rigid spacecraft and that arbitrary reorientation maneuvers of the spacecraft can be accomplished. Although the complete spacecraft equations are not controllable, the spacecraft equations are small time locally controllable in a reduced nonlinear sense. The reduced spacecraft dynamics cannot be asymptotically stabilized to any equilibrium attitude using a time-variant continuous feedback control law, but discontinuous feedback control strategies are constructed which stabilize any equilibrium attitude of the spacecraft in finite time. Consequently, reorientation of the spacecraft can be accomplished using discontinuous feedback control.
Krishnan, Hariharan↗
Spacecraft Attitude Tracking and Maneuver Using Combined Magnetic Actuators
A paper describes attitude-control algorithms using the combination of magnetic actuators with reaction wheel assemblies (RWAs) or other types of actuators such as thrusters. The combination of magnetic actuators with one or two RWAs aligned with different body axis expands the two-dimensional control torque to three-dimensional. The algorithms can guarantee the spacecraft attitude and rates to track the commanded attitude precisely. A design example is presented for nadir-pointing, pitch, and yaw maneuvers. The results show that precise attitude tracking can be reached and the attitude- control accuracy is comparable with RWA-based attitude control. When there are only one or two workable RWAs due to RWA failures, the attitude-control system can switch to the control algorithms for the combined magnetic actuators with the RWAs without going to the safe mode, and the control accuracy can be maintained. The attitude-control algorithms of the combined actuators are derived, which can guarantee the spacecraft attitude and rates to track the commanded values precisely. Results show that precise attitude tracking can be reached, and the attitude-control accuracy is comparable with 3-axis wheel control.
Zhou, Zhiqiang↗
Wild 2 approach maneuver strategy for Stardust spacecraft
14th AAS/AIAA Space Flight Mechanics Meeting Maui, Hawaii, USAStardust will return samples of dust from comet Wild 2 to be collected during an encounter in January 2004. Approach to Wild 2 will be performed with a number of trajectory correction maneuvers following a period of solar conjunction ending in early October 2003.
maneuver↗
Three axis rotational maneuver and vibration stabilization of elastic spacecraft
A control law for three-axis rotational maneuvers of a spacecraft beam-tip body configuration based on non-linear inversion and modal velocity feedback is presented. A decoupling attitude control law is presented such that in the closed-loop system the attitude angles of the spacecraft are independently controlled, using the control moments acting on the space vehicle. This controller asymptotically decouples the flexible dynamics from the rigid one and also allows the decomposition of the elastic dynamics into two subsystems representing the transverse deflections of the beam in two orthogonal planes. These low-order subsystems are used for the derivation of a modal velocity feedback stabilizer using the force and moment actuators at the end body. Simulation results are presented to show the capability of the controller.
Singh, Sahjendra N.↗
Orbit trim maneuver design and implementation for the 1975 Mars Viking Mission
The Viking Mission included the insertion of two unmanned spacecraft into orbit about Mars and the deployment of a soft-lander from each. A description is presented of the adaptive design and implementation of the spacecraft propulsive maneuvers as these flights progressed, taking into account also the inflight results for the orbital phase of the primary Viking mission. The design process included the selection of target parameters and the minimization of both propellant usage and the effects of execution errors, while complying with mission and operational constraints. All maneuvers performed to trim the spacecraft orbits are considered. Attention is given to navigation requirements, geometry definitions and terminology, maneuver mechanization and constraints, and maneuver capability.
Hintz, G. R.↗
Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation
Navigation of the orbit phase of the Near Earth steroid Rendezvous (NEAR) mission will re,quire determination of certain physical parameters describing the size, shape, gravity field, attitude and inertial properties of Eros. Prior to launch, little was known about Eros except for its orbit which could be determined with high precision from ground based telescope observations. Radar bounce and light curve data provided a rough estimate of Eros shape and a fairly good estimate of the pole, prime meridian and spin rate. However, the determination of the NEAR spacecraft orbit requires a high precision model of Eros's physical parameters and the ground based data provides only marginal a priori information. Eros is the principal source of perturbations of the spacecraft's trajectory and the principal source of data for determining the orbit. The initial orbit determination strategy is therefore concerned with developing a precise model of Eros. The original plan for Eros orbital operations was to execute a series of rendezvous burns beginning on December 20,1998 and insert into a close Eros orbit in January 1999. As a result of an unplanned termination of the rendezvous burn on December 20, 1998, the NEAR spacecraft continued on its high velocity approach trajectory and passed within 3900 km of Eros on December 23, 1998. The planned rendezvous burn was delayed until January 3, 1999 which resulted in the spacecraft being placed on a trajectory that slowly returns to Eros with a subsequent delay of close Eros orbital operations until February 2001. The flyby of Eros provided a brief glimpse and allowed for a crude estimate of the pole, prime meridian and mass of Eros. More importantly for navigation, orbit determination software was executed in the landmark tracking mode to determine the spacecraft orbit and a preliminary shape and landmark data base has been obtained. The flyby also provided an opportunity to test orbit determination operational procedures that will be used in February of 2001. The initial attitude and spin rate of Eros, as well as estimates of reference landmark locations, are obtained from images of the asteroid. These initial estimates are used as a priori values for a more precise refinement of these parameters by the orbit determination software which combines optical measurements with Doppler tracking data to obtain solutions for the required parameters. As the spacecraft is maneuvered; closer to the asteroid, estimates of spacecraft state, asteroid attitude, solar pressure, landmark locations and Eros physical parameters including mass, moments of inertia and gravity harmonics are determined with increasing precision. The determination of the elements of the inertia tensor of the asteroid is critical to spacecraft orbit determination and prediction of the asteroid attitude. The moments of inertia about the principal axes are also of scientific interest since they provide some insight into the internal mass distribution. Determination of the principal axes moments of inertia will depend on observing free precession in the asteroid's attitude dynamics. Gravity harmonics are in themselves of interest to science. When compared with the asteroid shape, some insight may be obtained into Eros' internal structure. The location of the center of mass derived from the first degree harmonic coefficients give a direct indication of overall mass distribution. The second degree harmonic coefficients relate to the radial distribution of mass. Higher degree harmonics may be compared with surface features to gain additional insight into mass distribution. In this paper, estimates of Eros physical parameters obtained from the December 23,1998 flyby will be presented. This new knowledge will be applied to simplification of Eros orbital operations in February of 2001. The resulting revision to the orbit determination strategy will also be discussed.
Miller, J. K.↗
Study to define logic associated with CMGS to maneuver and stabilize an orbiting spacecraft
A study was conducted to define the logic associated with the control moment gyroscopes to maneuver and stabilize an orbiting spacecraft. The study objectives are as follows: (1) to define mission requirements and feasible attitudes for a shuttle-like vehicle that will meet mission objectives, (2) to determine the control moment gyroscope (CMG) and system configurations that will best meet overall mission requirements, (3) to define all of the software required to manage and control the selected CMG systems, and (4) to verify by computer simulation the adequacy of the selected CMG system and specified software package in meeting the overall mission requirements.
Source record↗
Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the LADEE Spacecraft
This paper presents results from a contingency trajectory analysis performed for the Lunar Atmosphere Dust Environment Explorer (LADEE) mission in the event of a missed lunar-orbit insertion (LOI) maneuver by the LADEE spacecraft. The effects of varying solar perturbations in the vicinity of the weak stability boundary (WSB) in the Sun-Earth system on the trajectory design are analyzed and discussed. It is shown that geocentric recovery trajectory options existed for the LADEE spacecraft, depending on the spacecrafts recovery time to perform an Earth escape-prevention maneuver after the hypothetical LOI maneuver failure and subsequent path traveled through the Sun-Earth WSB. If Earth-escape occurred, a heliocentric recovery option existed, but with reduced science capability for the spacecraft in an eccentric, not circular near-equatorial retrograde lunar orbit.
Genova, Anthony L.↗
Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the LADEE Spacecraft
This paper presents results from a contingency trajectory analysis performed for the Lunar Atmosphere & Dust Environment Explorer (LADEE) mission in the event of a missed lunar-orbit insertion (LOI) maneuver by the LADEE spacecraft. The effects of varying solar perturbations in the vicinity of the weak stability boundary (WSB) in the Sun-Earth system on the trajectory design are analyzed and discussed. It is shown that geocentric recovery trajectory options existed for the LADEE spacecraft, depending on the spacecraft's recovery time to perform an Earth escape-prevention maneuver after the hypothetical LOI maneuver failure and subsequent path traveled through the Sun-Earth WSB. If Earth-escape occurred, a heliocentric recovery option existed, but with reduced science capacapability for the spacecraft in an eccentric, not circular near-equatorial retrograde lunar orbit.
LADEE↗
Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the LADEE Spacecraft
A contingency trajectory analysis was performed for NASA Ames Research Center’s (ARC’s) Lunar Atmosphere and Dust Environment Explorer (LADEE) spacecraft in case of a missed lunar orbit insertion (LOI) maneuver. Recovery trajectory options are shown to exist for all LADEE launch opportunities throughout a one year period. Recovery ΔV costs primarily depended on the spacecraft’s apogee location on or near the Sun-Earth weak stability boundary (WSB) and the time needed by the spacecraft to recover (e.g. to “wake up” from “safe” mode) to perform an escape prevention maneuver after the missed LOI.
Genova, A. L.↗