GEMINI RENDEZVOUS EVALUATION POD (REP) - ARTIST CONCEPT

NASA Image and Video Library

1965-08-01

S65-28653 (August 1965) --- Rendezvous Evaluation Pod (REP) in orbit is approached by Gemini spacecraft as seen in this artist's concept using an actual photograph taken on the Gemini-4 mission. The REP is superimposed over a Gemini-4 Earth-sky picture of cloud formations over an ocean. The REP will be used by the crew of the Gemini-5 spacecraft to practice rendezvous techniques.

SEP ENCKE-87 and Halley rendezvous studies and improved S/C model implementation in HILTOP

NASA Technical Reports Server (NTRS)

Horsewood, J. L.; Mann, F. I.

1978-01-01

Studies were conducted to determine the performance requirements for projected state-of-the-art SEP spacecrafts boosted by the Shuttle/IUS to perform a rendezvous with the comet Halley and a rendezvous with the comet Encke during its 1977 apparition. The spacecraft model of the standard HILTOP computer program was assumed. Numerical and graphical results summarizing the studies are presented.

Gemini rendezvous docking simulator

NASA Image and Video Library

1963-11-04

Multiple exposure of Gemini rendezvous docking simulator. Francis B. Smith wrote in his paper "Simulators for Manned Space Research," "The rendezvous and docking operation of the Gemini spacecraft with the Agena and of the Apollo Command Module with the Lunar Excursion Module have been the subject of simulator studies for several years. [This figure] illustrates the Gemini-Agena rendezvous docking simulator at Langley. The Gemini spacecraft was supported in a gimbal system by an overhead crane and gantry arrangement which provided 6 degrees of freedom - roll, pitch, yaw, and translation in any direction - all controllable by the astronaut in the spacecraft. Here again the controls fed into a computer which in turn provided an input to the servos driving the spacecraft so that it responded to control motions in a manner which accurately simulated the Gemini spacecraft." A.W. Vogeley further described the simulator in his paper "Discussion of Existing and Planned Simulators For Space Research," "Docking operations are considered to start when the pilot first can discern vehicle target size and aspect and terminate, of course, when soft contact is made. ... This facility enables simulation of the docking operation from a distance of 200 feet to actual contact with the target. A full-scale mock-up of the target vehicle is suspended near one end of the track. ... On [the Agena target] we have mounted the actual Agena docking mechanism and also various types of visual aids. We have been able to devise visual aids which have made it possible to accomplish nighttime docking with as much success as daytime docking." -- Published in Barton C. Hacker and James M. Grimwood, On the Shoulders of Titans: A History of Project Gemini, NASA SP-4203; Francis B. Smith, "Simulators for Manned Space Research," Paper presented at the 1966 IEEE International convention, March 21-25, 1966; A.W. Vogeley, "Discussion of Existing and Planned Simulators For Space Research," Paper presented at the Conference on the Role of Simulation in Space Technology, August 17-21, 1964.

KSC-04pd1595

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, an Orbital Sciences technician works with wiring on the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator, a spacecraft developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1592

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, Orbital Sciences workers remove the canister from the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator, a spacecraft developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1599

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, Orbital Sciences technicians watch closely as the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator is lowered onto a stand. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1594

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator is revealed after its protective cover has been removed. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1593

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator is revealed after its protective cover has been removed. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04PD-1593

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. At Vandenberg Air Force Base in California, the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator is revealed after its protective cover has been removed. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbitals Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

Servicing and Deployment of National Resources in Sun-Earth Libration Point Orbits

NASA Technical Reports Server (NTRS)

Folta, David C.; Beckman, Mark; Mar, Greg C.; Mesarch, Michael; Cooley, Steven; Leete, Steven J.

2002-01-01

Spacecraft travel between the Sun-Earth system, the Earth-Moon system, and beyond has received extensive attention recently. The existence of a connection between unstable regions enables mission designers to envision scenarios of multiple spacecraft traveling cheaply from system to system, rendezvousing, servicing, and refueling along the way. This paper presents examples of transfers between the Sun-Earth and Earth-Moon systems using a true ephemeris and perturbation model. It shows the (Delta)V costs associated with these transfers, including the costs to reach the staging region from the Earth. It explores both impulsive and low thrust transfer trajectories. Additionally, analysis that looks specifically at the use of nuclear power in libration point orbits and the issues associated with them such as inadvertent Earth return is addressed. Statistical analysis of Earth returns and the design of biased orbits to prevent any possible return are discussed. Lastly, the idea of rendezvous between spacecraft in libration point orbits using impulsive maneuvers is addressed.

GEMINI-TITAN (GT)-12 - EARTH SKY - AGENA RENDEZVOUS - OUTER SPACE

NASA Image and Video Library

1966-11-11

S66-62755 (11 Nov. 1966) --- Excellent stereo and side view of the Agena Target Docking Vehicle as seen from the Gemini-12 spacecraft during rendezvous and docking mission in space. The two spacecraft are 50 feet apart. Photo credit: NASA

Gemini Rendezvous Docking Simulator

NASA Image and Video Library

1964-05-11

Gemini Rendezvous Docking Simulator suspended from the roof of the Langley Research Center s aircraft hangar. Francis B. Smith wrote: The rendezvous and docking operation of the Gemini spacecraft with the Agena and of the Apollo Command Module with the Lunar Excursion Module have been the subject of simulator studies for several years. This figure illustrates the Gemini-Agena rendezvous docking simulator at Langley. The Gemini spacecraft was supported in a gimbal system by an overhead crane and gantry arrangement which provided 6 degrees of freedom - roll, pitch, yaw, and translation in any direction - all controllable by the astronaut in the spacecraft. Here again the controls fed into a computer which in turn provided an input to the servos driving the spacecraft so that it responded to control motions in a manner which accurately simulated the Gemini spacecraft. -- Published in Barton C. Hacker and James M. Grimwood, On the Shoulders of Titans: A History of Project Gemini, NASA SP-4203 Francis B. Smith, Simulators for Manned Space Research, Paper presented at the 1966 IEEE International convention, March 21-25, 1966.

Small Solar Electric Propulsion Spacecraft Concept for Near Earth Object and Inner Solar System Missions

NASA Technical Reports Server (NTRS)

Lang, Jared J.; Randolph, Thomas M.; McElrath, Timothy P.; Baker, John D.; Strange, Nathan J.; Landau, Damon; Wallace, Mark S.; Snyder, J. Steve; Piacentine, Jamie S.; Malone, Shane;

Six degree of freedom simulation system for evaluating automated rendezvous and docking spacecraft

NASA Technical Reports Server (NTRS)

Rourke, Kenneth H.; Tsugawa, Roy K.

1991-01-01

Future logistics supply and servicing vehicles such as cargo transfer vehicles (CTV) must have full 6 degree of freedom (6DOF) capability in order to perform requisite rendezvous, proximity operations, and capture operations. The design and performance issues encountered when developing a 6DOF maneuvering spacecraft are very complex with subtle interactions which are not immediately obvious or easily anticipated. In order to deal with these complexities and develop robust maneuvering spacecraft designs, a simulation system and associated family of tools are used at TRW for generating and validating spacecraft performance requirements and guidance algorithms. An overview of the simulator and tools is provided. These are used by TRW for autonomous rendezvous and docking research projects including CTV studies.

Rendezvous Integration Complexities of NASA Human Flight Vehicles

NASA Technical Reports Server (NTRS)

Brazzel, Jack P.; Goodman, John L.

2009-01-01

Propellant-optimal trajectories, relative sensors and navigation, and docking/capture mechanisms are rendezvous disciplines that receive much attention in the technical literature. However, other areas must be considered. These include absolute navigation, maneuver targeting, attitude control, power generation, software development and verification, redundancy management, thermal control, avionics integration, robotics, communications, lighting, human factors, crew timeline, procedure development, orbital debris risk mitigation, structures, plume impingement, logistics, and in some cases extravehicular activity. While current and future spaceflight programs will introduce new technologies and operations concepts, the complexity of integrating multiple systems on multiple spacecraft will remain. The systems integration task may become more difficult as increasingly complex software is used to meet current and future automation, autonomy, and robotic operation requirements.

Rendezvous and Proximity Operations of the Space Shuttle

NASA Technical Reports Server (NTRS)

Goodman, John L.

2005-01-01

Space Shuttle rendezous missions presented unique challenges that were not fully recognized when the Shuttle was designed. Rendezvous targets could be passive (i.e., no lights or transponders), and not designed to facilitate Shuttle rendezvous, proximity operations and retrieval. Shuttle reaction control system jet plume impingement on target spacecraft presented induced dynamics, structural loading and contamination concerns. These issues, along with limited forward reaction control system propellant, drove a change from the Gemimi/Apollo coelliptic profile heritage to a stable orbit profile, and the development of new proximity operations techniques. Multiple scientific and on-orbit servicing missions and crew exchange, assembly and replinishment flights to Mir and to the International Space Station drove further profile and piloting technique changes, including new relative navigation sensors and new computer generated piloting cues.

Low Cost Multiple Near Earth Object Missions

NASA Astrophysics Data System (ADS)

Smith, D. B.; Klaus, K.; Kaplan, M.

2009-12-01

Commercial spacecraft are available with efficient high power solar arrays and hybrid propulsion systems (Chemical and Solar Electric) that make possible multiple Near Earth Object Missions within Discovery budget limits. Our analysis is based on the Geosynchronous Transfer Orbit Capability (GTOC-3) solution. GTOC-3 assumptions: - Escape from Earth, rendezvous with 3 asteroids, then rendezvous with Earth - Departure velocity below 0.5 km/s - Launch between 2016 and 2025 - Total trip time less than 10 years - Minimum stay time of 60 days at each asteroid - Initial spacecraft mass of 2,000 kg - Thrust of 0.15 N and Isp of 3,000 s - Only Earth GAMs allowed (Rmin = 6,871 km) Preliminary results indicate that for mission objectives we can visit Apophis and any other 2 asteroids on this list or any other 3 asteroids listed. We have considered two spacecraft approaches to accomplish mission objectives: - Case 1: Chemical engine burn to the 1st target, and then solar electric to the 2nd and 3rd targets, or - Case 2: Solar electric propulsion to all 3 targets For both Cases, we assumed an instrument mass of up to 100 kg, power up to 100 W, and s/c bus pointing as good as 12 arc sec.Multi-NEO Mission Candidates

SEPS mission and system integration/interface requirements for the space transportation system. [Solar Electric Propulsion System

NASA Technical Reports Server (NTRS)

Cork, M. J.; Barnett, P. M.; Shaffer, J., Jr.; Doran, B. J.

1979-01-01

Earth escape mission requirements on Solar Electric Propulsion System (SEPS), and the interface definition and planned integration between SEPS, user spacecraft, and other elements of the STS. Emphasis is placed on the Comet rendezvous mission, scheduled to be the first SEPS user. Interactive SEPS interface characteristics with spacecraft and mission, as well as the multiple organizations and inter-related development schedules required to integrate the SEPS with spacecraft and STS, require early attention to definition of interfaces in order to assure a successful path to the first SEPS launch in July 1985

KSC-04pd1639

NASA Image and Video Library

2004-07-27

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft is on a work stand waiting for processing activities. The spacecraft was developed for NASA by Orbital Sciences Corporation in Dulles, Va., to prove technologies for locating and maneuvering near an orbiting satellite. DART will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1638

NASA Image and Video Library

2004-07-27

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft is placed on a work stand for processing activities. The spacecraft was developed for NASA by Orbital Sciences Corporation in Dulles, Va., to prove technologies for locating and maneuvering near an orbiting satellite. DART will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

Spacecraft Trajectory Analysis and Mission Planning Simulation (STAMPS) Software

NASA Technical Reports Server (NTRS)

Puckett, Nancy; Pettinger, Kris; Hallstrom,John; Brownfield, Dana; Blinn, Eric; Williams, Frank; Wiuff, Kelli; McCarty, Steve; Ramirez, Daniel; Lamotte, Nicole;

KSC-04pd1597

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, Orbital Sciences technicians check the bottom of the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator as it is raised of its platform. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1596

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, Orbital Sciences technicians check the bottom of the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator as it is raised off its platform. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1598

NASA Image and Video Library

2004-07-14

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, Orbital Sciences technicians observe closely the movement of the DART (Demonstration for Autonomous Rendezvous Technology) flight demonstrator as it is lowered onto a stand. The spacecraft was developed to prove technologies for locating and maneuvering near an orbiting satellite. Future applications of technologies developed by the DART project will benefit the nation in future space-vehicle systems development requiring in-space assembly, services or other autonomous rendezvous operations. Designed and developed for NASA by Orbital Sciences Corporation in Dulles, Va., the DART spacecraft will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

Mission requirements CSM-111/DM-2 Apollo/Soyuz test project

NASA Technical Reports Server (NTRS)

Blackmer, S. M.

1974-01-01

Test systems are developed for rendezvous and docking of manned spacecraft and stations that are suitable for use as a standard international system. This includes the rendezvous and docking of Apollo and Soyuz spacecraft, and crew transfer. The conduct of the mission will include: (1) testing of compatible rendezvous systems in orbit; (2) testing of universal docking assemblies; (3) verifying the techniques for transfer of cosmonauts and astronauts; (4) performing certain activities by U.S.A. and U.S.S.R. crews in joint flight; and (5) gaining of experience in conducting joint flights by U.S.A. and U.S.S.R. spacecraft, including, in case of necessity, rendering aid in emergency situations.

Cometary exploration in the shuttle era

NASA Technical Reports Server (NTRS)

Farquhar, R. W.; Wooden, W. H., II

1978-01-01

A comprehensive program plan for cometary exploration in the 1980-2000 time frame is proposed. Plans for ground-based observations, a Spacelab cometary observatory, and the Space Telescope are included in the observational program. The cometary mission sequence begins with a dual-spacecraft flyby of Halley's comet. The nominal mission strategy calls for a simultaneous launch of two spacecraft towards an intercept with Halley in March 1986. After the Halley encounter, the spacecraft are retargeted: one to intercept comet Borrelly in January 1988 and the other to intercept comet Tempel-2 in September 1988. The additional cometary intercepts are accomplished by utilizing a novel Earth-swingby technique. The next mission in the cometary program plan, a rendezvous with Encke's comet, is scheduled for launch in early 1990. It is planned to rendezvous with Encke in September 1992 at a heliocentric distance of 4 AU. Following this near-aphelion rendezvous, the spacecraft will remain with with Encke through the next two perihelion passages in February 1994 and May 1997. The rendezvous mission will be terminated about seven months after the second perihelion passage.

Short rendezvous missions for advanced Russian human spacecraft

NASA Astrophysics Data System (ADS)

Murtazin, Rafail F.; Budylov, Sergey G.

2010-10-01

The two-day stay of crew in a limited inhabited volume of the Soyuz-TMA spacecraft till docking to ISS is one of the most stressful parts of space flight. In this paper a number of possible ways to reduce the duration of the free flight phase are considered. The duration is defined by phasing strategy that is necessary for reduction of the phase angle between the chaser and target spacecraft. Some short phasing strategies could be developed. The use of such strategies creates more comfortable flight conditions for crew thanks to short duration and additionally it allows saving spacecraft's life support resources. The transition from the methods of direct spacecraft rendezvous using one orbit phasing (first flights of " Vostok" and " Soyuz" vehicles) to the currently used methods of two-day rendezvous mission can be observed in the history of Soviet manned space program. For an advanced Russian human rated spacecraft the short phasing strategy is recommended, which can be considered as a combination between the direct and two-day rendezvous missions. The following state of the art technologies are assumed available: onboard accurate navigation; onboard computations of phasing maneuvers; launch vehicle with high accuracy injection orbit, etc. Some operational requirements and constraints for the strategies are briefly discussed. In order to provide acceptable phase angles for possible launch dates the experience of the ISS altitude profile control can be used. As examples of the short phasing strategies, the following rendezvous missions are considered: direct ascent, short mission with the phasing during 3-7 orbits depending on the launch date (nominal or backup). For each option statistical modeling of the rendezvous mission is fulfilled, as well as an admissible phase angle range, accuracy of target state vector and addition fuel consumption coming out of emergency is defined. In this paper an estimation of pros and cons of all options is conducted.

Biobjective planning of GEO debris removal mission with multiple servicing spacecrafts

NASA Astrophysics Data System (ADS)

Jing, Yu; Chen, Xiao-qian; Chen, Li-hu

2014-12-01

The mission planning of GEO debris removal with multiple servicing spacecrafts (SScs) is studied in this paper. Specifically, the SScs are considered to be initially on the GEO belt, and they should rendezvous with debris of different orbital slots and different inclinations, remove them to the graveyard orbit and finally return to their initial locations. Three key problems should be resolved here: task assignment, mission sequence planning and transfer trajectory optimization for each SSc. The minimum-cost, two-impulse phasing maneuver is used for each rendezvous. The objective is to find a set of optimal planning schemes with minimum fuel cost and travel duration. Considering this mission as a hybrid optimal control problem, a mathematical model is proposed. A modified multi-objective particle swarm optimization is employed to address the model. Numerous examples are carried out to demonstrate the effectiveness of the model and solution method. In this paper, single-SSc and multiple-SSc scenarios with the same amount of fuel are compared. Numerous experiments indicate that for a definite GEO debris removal mission, that which alternative (single-SSc or multiple-SSc) is better (cost less fuel and consume less travel time) is determined by many factors. Although in some cases, multiple-SSc scenarios may perform worse than single-SSc scenarios, the extra costs are considered worth the gain in mission safety and robustness.

Shared control on lunar spacecraft teleoperation rendezvous operations with large time delay

NASA Astrophysics Data System (ADS)

Ya-kun, Zhang; Hai-yang, Li; Rui-xue, Huang; Jiang-hui, Liu

2017-08-01

Teleoperation could be used in space on-orbit serving missions, such as object deorbits, spacecraft approaches, and automatic rendezvous and docking back-up systems. Teleoperation rendezvous and docking in lunar orbit may encounter bottlenecks for the inherent time delay in the communication link and the limited measurement accuracy of sensors. Moreover, human intervention is unsuitable in view of the partial communication coverage problem. To solve these problems, a shared control strategy for teleoperation rendezvous and docking is detailed. The control authority in lunar orbital maneuvers that involves two spacecraft as rendezvous and docking in the final phase was discussed in this paper. The predictive display model based on the relative dynamic equations is established to overcome the influence of the large time delay in communication link. We discuss and attempt to prove via consistent, ground-based simulations the relative merits of fully autonomous control mode (i.e., onboard computer-based), fully manual control (i.e., human-driven at the ground station) and shared control mode. The simulation experiments were conducted on the nine-degrees-of-freedom teleoperation rendezvous and docking simulation platform. Simulation results indicated that the shared control methods can overcome the influence of time delay effects. In addition, the docking success probability of shared control method was enhanced compared with automatic and manual modes.

Low-Cost Innovation in Spaceflight: The Near Earth Asteroid Rendezvous (NEAR) Shoemaker Mission

NASA Technical Reports Server (NTRS)

McCurdy, Howard E.

2005-01-01

On a spring day in 1996, at their research center in the Maryland countryside, representatives from the Johns Hopkins University Applied Physics Laboratory (APL) presented Administrator Daniel S. Goldin of the National Aeronautics and Space Administration (NASA) with a check for $3.6 million. 1 Two and a half years earlier, APL officials had agreed to develop a spacecraft capable of conducting an asteroid rendezvous and to do so for slightly more than $122 million. This was a remarkably low sum for a spacecraft due to conduct a planetaryclass mission. By contrast, the Mars Observer spacecraft launched in 1992 for an orbital rendezvous with the red planet had cost $479 million to develop, while the upcoming Cassini mission to Saturn required a spacecraft whose total cost was approaching $1.4 billion. In an Agency accustomed to cost overruns on major missions, the promise to build a planetary-class spacecraft for about $100 million seemed excessively optimistic.

Overview: Solar Electric Propulsion Concept Designs for SEP Technology Demonstration Mission

NASA Technical Reports Server (NTRS)

Mcguire, Melissa L.; Hack, Kurt J.; Manzella, David; Herman, Daniel

2014-01-01

JPC presentation of the Concept designs for NASA Solar Electric Propulsion Technology Demonstration mission paper. Multiple Solar Electric Propulsion Technology Demonstration Missions were developed to assess vehicle performance and estimated mission cost. Concepts ranged from a 10,000 kg spacecraft capable of delivering 4000 kg of payload to one of the Earth Moon Lagrange points in support of future human-crewed outposts to a 180 kg spacecraft capable of performing an asteroid rendezvous mission after launched to a geostationary transfer orbit as a secondary payload.

Optimal cooperative time-fixed impulsive rendezvous

NASA Technical Reports Server (NTRS)

Mirfakhraie, Koorosh; Conway, Bruce A.

1990-01-01

New capabilities have been added to a method that had been developed for determining optimal, i.e., minimum fuel, trajectories for the fixed-time cooperative rendezvous of two spacecraft. The method utilizes the primer vector theory. The new capabilities enable the method to accomodate cases in which there are fuel constraints on the spacecraft and/or enable the addition of a mid-course impulse to one of the vehicle's trajectories. Results are presented for a large number of cases, and the effect of varying parameters, such as vehicle fuel constraints, vehicle initial masses, and time allowed for the rendezvous, is demonstrated.

Trajectory design for an ion drive asteroids rendezvous mission launched into an Ariane geostationary transfer orbit

NASA Astrophysics Data System (ADS)

Pietrass, A. E.

1984-08-01

AMSAT has conceived an asteroid rendezvous mission which would consist of an Ariane-launched, 3-axis-stabilized, 350-kg spacecraft utilizing both mercury and solar electric ion propulsion. The spacecraft is to be equipped with a science instrument platform with a mass of approximately 30 to 50 kg. Practically uninterrupted earth departure opportunities are found for targets such as 4 Vesta, 8 Flora, and 19 Fortuna from 1986 through 1988. The 7 to 8 year mission would allow for a second rendezvous of 4 Vesta, and marginal additional fuel would make close flybys of targets feasible. Through the use of parameter optimization techniques, trajectories can be generated and the inclusion of constraints due to spacecraft techology, tour design, and navigation can be facilitated.

14 CFR 1214.111 - Rendezvous services.

Code of Federal Regulations, 2011 CFR

2011-01-01

... Earth of the orbiting spacecraft (or part thereof), including a spacecraft deployed earlier on the same... orbiting spacecraft to Earth. (3) Revisit of an orbiting spacecraft for purposes such as resupply, repair...

KSC-04pd1636

NASA Image and Video Library

2004-07-27

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft is raised to a vertical position. It will be lifted onto a test stand for launch processing activities. The spacecraft was developed for NASA by Orbital Sciences Corporation in Dulles, Va., to prove technologies for locating and maneuvering near an orbiting satellite. DART will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

KSC-04pd1637

NASA Image and Video Library

2004-07-27

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft is raised to a vertical position. It will be lifted onto a test stand for launch processing activities. The spacecraft was developed for NASA by Orbital Sciences Corporation in Dulles, Va., to prove technologies for locating and maneuvering near an orbiting satellite. DART will be launched on a Pegasus launch vehicle. At about 40,000 feet over the Pacific Ocean, the Pegasus will be released from Orbital’s Stargazer L-1011 aircraft, fire its rocket motors and boost DART into a polar orbit approximately 472 miles by 479 miles. Once in orbit, DART will rendezvous with a target satellite, the Multiple Paths, Beyond-Line-of-Site Communications satellite, also built by Orbital Sciences. DART will then perform several close proximity operations, such as moving toward and away from the satellite using navigation data provided by onboard sensors. DART is scheduled for launch no earlier than Oct. 18.

Rendezvous missions to temporarily captured near Earth asteroids

NASA Astrophysics Data System (ADS)

Brelsford, S.; Chyba, M.; Haberkorn, T.; Patterson, G.

2016-04-01

Missions to rendezvous with or capture an asteroid present significant interest both from a geophysical and safety point of view. They are key to the understanding of our solar system and are stepping stones for interplanetary human flight. In this paper, we focus on a rendezvous mission with 2006 RH120, an asteroid classified as a Temporarily Captured Orbiter (TCO). TCOs form a new population of near Earth objects presenting many advantages toward that goal. Prior to the mission, we consider the spacecraft hibernating on a Halo orbit around the Earth-Moon's L2 libration point. The objective is to design a transfer for the spacecraft from the parking orbit to rendezvous with 2006 RH120 while minimizing the fuel consumption. Our transfers use indirect methods, based on the Pontryagin Maximum Principle, combined with continuation techniques and a direct method to address the sensitivity of the initialization. We demonstrate that a rendezvous mission with 2006 RH120 can be accomplished with low delta-v. This exploratory work can be seen as a first step to identify good candidates for a rendezvous on a given TCO trajectory.

Automatic trajectory planning for low-thrust active removal mission in low-earth orbit

NASA Astrophysics Data System (ADS)

Di Carlo, Marilena; Romero Martin, Juan Manuel; Vasile, Massimiliano

2017-03-01

In this paper two strategies are proposed to de-orbit up to 10 non-cooperative objects per year from the region within 800 and 1400 km altitude in Low Earth Orbit (LEO). The underlying idea is to use a single servicing spacecraft to de-orbit several objects applying two different approaches. The first strategy is analogous to the Traveling Salesman Problem: the servicing spacecraft rendezvous with multiple objects in order to physically attach a de-orbiting kit that reduces the perigee of the orbit. The second strategy is analogous to the Vehicle Routing Problem: the servicing spacecraft rendezvous and docks with an object, spirals it down to a lower altitude orbit, undocks, and then spirals up to the next target. In order to maximise the number of de-orbited objects with minimum propellant consumption, an optimal sequence of targets is identified using a bio-inspired incremental automatic planning and scheduling discrete optimisation algorithm. The optimisation of the resulting sequence is realised using a direct transcription method based on an asymptotic analytical solution of the perturbed Keplerian motion. The analytical model takes into account the perturbations deriving from the J2 gravitational effect and the atmospheric drag.

Orbit Modification of Earth-Crossing Asteroids/Comets Using Rendezvous Spacecraft and Laser Ablation

NASA Technical Reports Server (NTRS)

Park, Sang-Young; Mazanek, Daniel D.

2005-01-01

This report describes the approach and results of an end-to-end simulation to deflect a long-period comet (LPC) by using a rapid rendezvous spacecraft and laser ablation system. The laser energy required for providing sufficient deflection DELTA V and an analysis of possible intercept/rendezvous spacecraft trajectories are studied in this analysis. These problems minimize a weighted sum of the flight time and required propellant by using an advanced propulsion system. The optimal thrust-vector history and propellant mass to use are found in order to transfer a spacecraft from the Earth to a targeted celestial object. One goal of this analysis is to formulate an optimization problem for intercept/rendezvous spacecraft trajectories. One approach to alter the trajectory of the object in a highly controlled manner is to use pulsed laser ablative propulsion. A sufficiently intense laser pulse ablates the surface of a near-Earth object (NEO) by causing plasma blowoff. The momentum change from a single laser pulse is very small. However, the cumulative effect is very effective because the laser can interact with the object over long periods of time. The laser ablation technique can overcome the mass penalties associated with other nondisruptive approaches because no propellant is required to generate the DELTA V (the material of the celestial object is the propellant source). Additionally, laser ablation is effective against a wide range of surface materials and does not require any landing or physical attachment to the object. For diverting distant asteroids and comets, the power and optical requirements of a laser ablation system on or near the Earth may be too extreme to contemplate in the next few decades. A hybrid solution would be for a spacecraft to carry a laser as a payload to a particular celestial body. The spacecraft would require an advanced propulsion system capable of rapid rendezvous with the object and an extremely powerful electrical generator, which is likely needed for the propulsion system as well. The spacecraft would station-keep with the object at a small standoff distance while the laser ablation is performed.

Concurrent image-based visual servoing with adaptive zooming for non-cooperative rendezvous maneuvers

NASA Astrophysics Data System (ADS)

Pomares, Jorge; Felicetti, Leonard; Pérez, Javier; Emami, M. Reza

2018-02-01

An image-based servo controller for the guidance of a spacecraft during non-cooperative rendezvous is presented in this paper. The controller directly utilizes the visual features from image frames of a target spacecraft for computing both attitude and orbital maneuvers concurrently. The utilization of adaptive optics, such as zooming cameras, is also addressed through developing an invariant-image servo controller. The controller allows for performing rendezvous maneuvers independently from the adjustments of the camera focal length, improving the performance and versatility of maneuvers. The stability of the proposed control scheme is proven analytically in the invariant space, and its viability is explored through numerical simulations.

Attitude control challenges for earth orbiters of the 1980's

NASA Technical Reports Server (NTRS)

Hibbard, W.

1980-01-01

Experience gained in designing attitude control systems for orbiting spacecraft of the late 1980's is related. Implications for satellite attitude control design of the guidance capabilities, rendezvous and recovery requirements, use of multiple-use spacecraft and the development of large spacecraft associated with the advent of the Space Shuttle are considered. Attention is then given to satellite attitude control requirements posed by the Tracking and Data Relay Satellite System, the Global Positioning System, the NASA End-to-End Data System, and Shuttle-associated subsatellites. The anticipated completion and launch of the Space Telescope, which will provide one of the first experiences with the new generation of attitude control, is also pointed out.

Automated Rendezvous and Docking Sensor Testing at the Flight Robotics Laboratory

NASA Technical Reports Server (NTRS)

Howard, Richard T.; Williamson, Marlin L.; Johnston, Albert S.; Brewster, Linda L.; Mitchell, Jennifer D.; Cryan, Scott P.; Strack, David; Key, Kevin

2007-01-01

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as Automated Rendezvous and Docking, (AR&D).) The crewed versions of the spacecraft may also perform AR&D, possibly with a different level of automation and/or autonomy, and must also provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success of the Exploration Program. NASA has the responsibility to determine whether the Crew Exploration Vehicle (CEV) contractor-proposed relative navigation sensor suite will meet the CEV requirements. The relatively low technology readiness of relative navigation sensors for AR&D has been carried as one of the CEV Projects top risks. The AR&D Sensor Technology Project seeks to reduce this risk by increasing technology maturation of selected relative navigation sensor technologies through testing and simulation, and to allow the CEV Project to assess the relative navigation sensors.

Autonomous spacecraft rendezvous and docking

NASA Technical Reports Server (NTRS)

Tietz, J. C.; Almand, B. J.

1985-01-01

A storyboard display is presented which summarizes work done recently in design and simulation of autonomous video rendezvous and docking systems for spacecraft. This display includes: photographs of the simulation hardware, plots of chase vehicle trajectories from simulations, pictures of the docking aid including image processing interpretations, and drawings of the control system strategy. Viewgraph-style sheets on the display bulletin board summarize the simulation objectives, benefits, special considerations, approach, and results.

Autonomous spacecraft rendezvous and docking

NASA Astrophysics Data System (ADS)

Tietz, J. C.; Almand, B. J.

A storyboard display is presented which summarizes work done recently in design and simulation of autonomous video rendezvous and docking systems for spacecraft. This display includes: photographs of the simulation hardware, plots of chase vehicle trajectories from simulations, pictures of the docking aid including image processing interpretations, and drawings of the control system strategy. Viewgraph-style sheets on the display bulletin board summarize the simulation objectives, benefits, special considerations, approach, and results.

Multiple NEO Rendezvous, Reconnaissance and In Situ Exploration

NASA Astrophysics Data System (ADS)

Klaus, K.; Elsperman, M. S.; Cook, T.; Smith, D.

2010-12-01

We propose a two spacecraft mission (Mother Ship and Small Body Lander) rendezvous with multiple Near Earth Objects (NEO). This two spacecraft mission mimics the likely architecture approach that human explorers will use: a “mother ship”(MS) designed to get from Earth to the NEO and a “Small Body Lander”(SBL) that performs in situ investigation on or close to the NEO’s surface. The MS carries the SBL to the target NEO. Once at the target NEO, the MS conducts an initial reconnaissance in order to produce a high resolution map of the surface. This map is used to identify coordinates of interest which are sent to the SBL. The SBL un-docks from the MS to rendezvous with the NEO and collect data. Landings are possible, though the challenges of anchoring to the NEO surface are significant. The SBL design is flexible and adaptable, enabling science data collection on or near the surface. After surface investigations are completed on the first NEO, the SBL will return and autonomously rendezvous and dock with the MS. The MS then goes to the next NEO target. During transit to the next NEO, the SBL could be refueled by the MS, a TRL8 capability demonstrated on the DARPA/NASA Orbital Express mission in 2007, or alternately sized to operate without requiring refueling depending on the mission profile. The mission goals are to identify surface hazards; quantify engineering boundary conditions for future human visits, and identify resources for future exploitation. The mission goals will be accomplished through the execution of key mission objectives: (1) high-resolution surface topography; (2) surface composition and mineralogy; (3) radiation environment near NEO; and (4) mechanical properties of the surface. Essential SBL instruments include: a) LIDAR (Obj. 1); b) 3D, high- resolution hyperspectral imaging cameras (Obj. 2); c) radiation sensor package (Obj. 3); and d) strain gauges (Obj. 4). Additional or alternative instruments could include: e) x-ray fluorescence or laser-induced breakdown spectroscopy (LIBS) sensor package (Obj. 2); f) gamma ray/neutron spectrometry package (Obj. 2); and g) radiometer package (to address variations in thermal environment). The ability to reach, survey, sample, and analyze multiple NEOs at close proximity is an enormous capability that can enable NASA to rapidly achieve the primary Exploration Precursor Robotic Mission (xPRM) Program goal of characterizing NEOs for future human exploration. Instead of launching multiple dedicated missions to each NEO of interest, a multi-NEO sortie mission can be planned and executed to achieve the same mission objectives with one launch, dramatically reducing the cost of NEO exploration. Collectively, our NEO Exploration System Architecture provides solutions for a wide variety of exploration activities using a common spacecraft bus and common core instrumentation for the spacecraft. This engineering consistency will substantially improve the probability of mission success, increase the likelihood of maintaining an aggressive launch schedule, and decrease the total cost of multiple missions. NASA successfully used this approach with the robotic precursors leading up to the Apollo missions, and we see significant benefits from this same programmatic approach for the xPRM program.

Cost-Effective NEO Characterization Using Solar Electric Propulsion (SEP)

NASA Astrophysics Data System (ADS)

Dissly, R. W.; Reinert, R.; Mitchell, S.

2003-05-01

We present a cost-effective multiple NEO rendezvous mission design optimized around the capabilities of Ball's 200-kg NEOX Solar Electric Propelled microsatellite. The NEOX spacecraft is 3-axis stabilized with better-than 1 milliradian pointing accuracy to serve as an excellent imaging platform; its DSN compatible telecommunications subsystem can support a 6.4-kbps downlink rate at 3 AU earth range. The spacecraft mass is <200kg at launch to allow launch as a cost-effective secondary payload. It uses proven SEP technology to provide 12km/s of Delta-V, which enables multiple rendezvous' in a single mission. Cost-effectiveness is optimized by launch as a secondary payload (e.g., Ariane-5 ASAP) or as a multiple manifest on a single dedicated launch vehicle (e.g., 4 on a Delta-II 2925). Following separation from the LV, we describe a candidate mission profile that minimizes cost by using the spacecraft's 12km/s of SEP Delta-V to allow orbiting up to 4 separate NEO's. Orbiting as opposed to flying by augments the mission's science return by providing the NEO mass and by allowing multiple phase angle imaging. The NEOX Spacecraft has the capability to support a 20kg payload drawing 100W average during SEP cruise, with >1kW available during the NEO orbital phase when the SEP thrusters are not powered. We will present a candidate payload suite that includes a visible/NIR imager, a laser altimeter, and a set of small, self-righting surface probes that can be used to assess the geophysical state of the object surface and near-surface environments. The surface probe payload notionally includes a set of cameras for imaging the body surface at mm-scale resolution, an accelerometer package to measure surface mechanical properties upon probe impact, a Langmuir probe to measure the electrostatic gradient immediately above the object surface, and an explosive charge that can be remotely detonated at the end of the surface mission to excavate an artificial crater that can be remotely observed from the orbiting spacecraft.

Autonomous Onboard Science Data Analysis for Comet Missions

NASA Technical Reports Server (NTRS)

Thompson, David R.; Tran, Daniel Q.; McLaren, David; Chien, Steve A.; Bergman, Larry; Castano, Rebecca; Doyle, Richard; Estlin, Tara; Lenda, Matthew

2012-01-01

Coming years will bring several comet rendezvous missions. The Rosetta spacecraft arrives at Comet 67P/Churyumov-Gerasimenko in 2014. Subsequent rendezvous might include a mission such as the proposed Comet Hopper with multiple surface landings, as well as Comet Nucleus Sample Return (CNSR) and Coma Rendezvous and Sample Return (CRSR). These encounters will begin to shed light on a population that, despite several previous flybys, remains mysterious and poorly understood. Scientists still have little direct knowledge of interactions between the nucleus and coma, their variation across different comets or their evolution over time. Activity may change on short timescales so it is challenging to characterize with scripted data acquisition. Here we investigate automatic onboard image analysis that could act faster than round-trip light time to capture unexpected outbursts and plume activity. We describe one edge-based method for detect comet nuclei and plumes, and test the approach on an existing catalog of comet images. Finally, we quantify benefits to specific measurement objectives by simulating a basic plume monitoring campaign.

Robust H ∞ Control for Spacecraft Rendezvous with a Noncooperative Target

PubMed Central

Wu, Shu-Nan; Zhou, Wen-Ya; Tan, Shu-Jun; Wu, Guo-Qiang

2013-01-01

The robust H ∞ control for spacecraft rendezvous with a noncooperative target is addressed in this paper. The relative motion of chaser and noncooperative target is firstly modeled as the uncertain system, which contains uncertain orbit parameter and mass. Then the H ∞ performance and finite time performance are proposed, and a robust H ∞ controller is developed to drive the chaser to rendezvous with the non-cooperative target in the presence of control input saturation, measurement error, and thrust error. The linear matrix inequality technology is used to derive the sufficient condition of the proposed controller. An illustrative example is finally provided to demonstrate the effectiveness of the controller. PMID:24027446

Designing the STS-134 Re-Rendezvous: A Preparation for Future Crewed Rendezvous Missions

NASA Technical Reports Server (NTRS)

Stuit, Timothy D.

2011-01-01

In preparation to provide the capability for the Orion spacecraft, also known as the Multi-Purpose Crew Vehicle (MPCV), to rendezvous with the International Space Station (ISS) and future spacecraft, a new suite of relative navigation sensors are in development and were tested on one of the final Space Shuttle missions to ISS. The National Aeronautics and Space Administration (NASA) commissioned a flight test of prototypes of the Orion relative navigation sensors on STS-134, in order to test their performance in the space environment during the nominal rendezvous and docking, as well as a re-rendezvous dedicated to testing the prototype sensors following the undocking of the Space Shuttle orbiter at the end of the mission. Unlike the rendezvous and docking at the beginning of the mission, the re-rendezvous profile replicates the newly designed Orion coelliptic approach trajectory, something never before attempted with the shuttle orbiter. Therefore, there were a number of new parameters that needed to be conceived of, designed, and tested for this rerendezvous to make the flight test successful. Additionally, all of this work had to be integrated with the normal operations of the ISS and shuttle and had to conform to the constraints of the mission and vehicles. The result of this work is a separation and rerendezvous trajectory design that would not only prove the design of the relative navigation sensors for the Orion vehicle, but also would serve as a proof of concept for the Orion rendezvous trajectory itself. This document presents the analysis and decision making process involved in attaining the final STS-134 re-rendezvous design.

14 CFR § 1214.111 - Rendezvous services.

Code of Federal Regulations, 2014 CFR

2014-01-01

... Earth of the orbiting spacecraft (or part thereof), including a spacecraft deployed earlier on the same Space Shuttle flight. (2) Exchange of a spacecraft (or part thereof) delivered to orbit on a particular... orbiting spacecraft to Earth. (3) Revisit of an orbiting spacecraft for purposes such as resupply, repair...

Searching for organics on the dwarf planet Ceres

NASA Astrophysics Data System (ADS)

Nayak, Michael

The Herschel Space Observatory recently detected the presence of water vapor in observations of Ceres, bringing it into the crosshairs of the search for the building blocks of life in the solar system. I present a mission concept designed in collaboration with the NASA Ames Research Center for a two-probe mission to the dwarf planet Ceres, utilizing a pair of small low-cost spacecraft. The primary spacecraft will carry both a mass and an infrared spectrometer to characterize the detected vapor. Shortly after its arrival a second and largely similar spacecraft will impact Ceres to create an impact ejecta "plume" timed to enable a rendezvous and sampling by the primary spacecraft. This enables additional subsurface chemistry, volatile content and material characterization, and new science complementary to the Dawn spacecraft, the first to arrive at Ceres. Science requirements, candidate instruments, rendezvous trajectories, spacecraft design and comparison with Dawn science are detailed.

Near Earth asteroid rendezvous

NASA Technical Reports Server (NTRS)

1992-01-01

The Spacecraft Design Course is the capstone design class for the M.S. in astronautics at the Naval Postgraduate School. The Fall 92 class designed a spacecraft for the Near Earth Asteroid Rendezvous Mission (NEAR). The NEAR mission uses a robotic spacecraft to conduct up-close reconnaissance of a near-earth asteroid. Such a mission will provide information on Solar System formation and possible space resources. The spacecraft is intended to complete a NEAR mission as a relatively low-budget program while striving to gather as much information about the target asteroid as possible. A complete mission analysis and detailed spacecraft design were completed. Mission analysis includes orbit comparison and selection, payload and telemetry requirements, spacecraft configuration, and launch vehicle selection. Spacecraft design includes all major subsystems: structure, electrical power, attitude control, propulsion, payload integration, and thermal control. The resulting spacecraft demonstrates the possibility to meet the NEAR mission requirements using existing technology, 'off-the-shelf' components, and a relatively low-cost launch vehicle.

Orbit Maintenance and Navigation of Human Spacecraft at Cislunar Near Rectilinear Halo Orbits

NASA Technical Reports Server (NTRS)

Davis, Diane; Bhatt, Sagar; Howell, Kathleen; Jang, Jiann-Woei; Whitley, Ryan; Clark, Fred; Guzzetti, Davide; Zimovan, Emily; Barton, Gregg

2017-01-01

Multiple studies have concluded that Earth-Moon libration point orbits are attractive candidates for staging operations. The Near Rectilinear Halo Orbit (NRHO), a member of the Earth-Moon halo orbit family, has been singularly demonstrated to meet multi-mission architectural constraints. In this paper, the challenges associated with operating human spacecraft in the NRHO are evaluated. Navigation accuracies and human vehicle process noise effects are applied to various station keeping strategies in order to obtain a reliable orbit maintenance algorithm. Additionally, the ability to absorb missed burns, construct phasing maneuvers to avoid eclipses and conduct rendezvous and proximity operations are examined.

Reference equations of motion for automatic rendezvous and capture

NASA Technical Reports Server (NTRS)

Henderson, David M.

1992-01-01

The analysis presented in this paper defines the reference coordinate frames, equations of motion, and control parameters necessary to model the relative motion and attitude of spacecraft in close proximity with another space system during the Automatic Rendezvous and Capture phase of an on-orbit operation. The relative docking port target position vector and the attitude control matrix are defined based upon an arbitrary spacecraft design. These translation and rotation control parameters could be used to drive the error signal input to the vehicle flight control system. Measurements for these control parameters would become the bases for an autopilot or feedback control system (FCS) design for a specific spacecraft.

Usage of pre-flight data in short rendezvous mission of Soyuz-TMA spacecrafts

NASA Astrophysics Data System (ADS)

Murtazin, Rafail; Petrov, Nikolay

2014-01-01

The paper describes the reduction of the vehicle autonomous flight duration before docking to the ISS. The Russian Soyuz-TMA spacecraft dock to the ISS two days after launch. Due to the limited volume inside Soyuz-TMA the reduction of time until docking to the ISS is very important, since the long stay of the cosmonauts in the limited volume adds to the strain of the space flight. In the previous papers of the authors it was shown that the existing capabilities of Soyuz-TMA, the ISS and the ground control loop make it possible to transfer to the five-orbit rendezvous profile. However, the analysis of the cosmonauts' schedule on the launch day shows that its duration is at the allowable limit and that is why it is necessary to find a way to further reduce the flight duration of Soyuz-TMA before docking to less than five orbits. In a traditional rendezvous profile, the calculation of rendezvous burns begins only after determination of the actual vehicle insertion orbit. The paper describes an approach in which the first two rendezvous burns are performed as soon as the spacecraft reaches the reference orbit and the values of the burns are calculated prior to the launch based on the pre-flight data for the nominal insertion. This approach decreases the duration of the rendezvous by one orbit. The demonstration flight of a Progress vehicle using the proposed profile was implemented on August 1, 2012 and completely confirmed the correctness of the imbedded principles. The paper considers the possible improvements of the proposed approach and recovery from the contingencies.

The Advanced Video Guidance Sensor: Orbital Express and the Next Generation

NASA Technical Reports Server (NTRS)

Howard, Richard T.; Heaton, Andrew F.; Pinson, Robin M.; Carrington, Connie L.; Lee, James E.; Bryan, Thomas C.; Robertson, Bryan A.; Spencer, Susan H.; Johnson, Jimmie E.

2008-01-01

The Orbital Express (OE) mission performed the first autonomous rendezvous and docking in the history of the United States on May 5-6, 2007 with the Advanced Video Guidance Sensor (AVGS) acting as one of the primary docking sensors. Since that event, the OE spacecraft performed four more rendezvous and docking maneuvers, each time using the AVGS as one of the docking sensors. The Marshall Space Flight Center's (MSFC's) AVGS is a nearfield proximity operations sensor that was integrated into the Autonomous Rendezvous and Capture Sensor System (ARCSS) on OE. The ARCSS provided the relative state knowledge to allow the OE spacecraft to rendezvous and dock. The AVGS is a mature sensor technology designed to support Automated Rendezvous and Docking (AR&D) operations. It is a video-based laser-illuminated sensor that can determine the relative position and attitude between itself and its target. Due to parts obsolescence, the AVGS that was flown on OE can no longer be manufactured. MSFC has been working on the next generation of AVGS for application to future Constellation missions. This paper provides an overview of the performance of the AVGS on Orbital Express and discusses the work on the Next Generation AVGS (NGAVGS).

Agile Science Operations: A New Approach for Primitive Exploration Bodies

NASA Technical Reports Server (NTRS)

Chien, Steve A.; Thompson, David R.; Castillo-Rogez, Julie C.; Doyle, Richard; Estlin, Tara; Mclaren, David

2012-01-01

Primitive body exploration missions such as potential Comet Surface Sample Return or Trojan Tour and Rendezvous would challenge traditional operations practices. Earth-based observations would provide only basic understanding before arrival and many science goals would be defined during the initial rendezvous. It could be necessary to revise trajectories and observation plans to quickly characterize the target for safe, effective observations. Detection of outgassing activity and monitoring of comet surface activity are even more time constrained, with events occurring faster than round-trip light time. "Agile science operations" address these challenges with contingency plans that recognize the intrinsic uncertainty in the operating environment and science objectives. Planning for multiple alternatives can significantly improve the time required to repair and validate spacecraft command sequences. When appropriate, time-critical decisions can be automated and shifted to the spacecraft for immediate access to instrument data. Mirrored planning systems on both sides of the light-time gap permit transfer of authority back and forth as needed. We survey relevant science objectives, identifying time bottlenecks and the techniques that could be used to speed missions' reaction to new science data. Finally, we discuss the results of a trade study simulating agile observations during flyby and comet rendezvous scenarios. These experiments quantify instrument coverage of key surface features as a function of planning turnaround time. Careful application of agile operations techniques can play a significant role in realizing the Decadal Survey plan for primitive body exploration

Use of automated rendezvous trajectory planning to improve spacecraft operations efficiency

NASA Technical Reports Server (NTRS)

Mulder, Tom A.

1991-01-01

The current planning process for space shuttle rendezvous with a second Earth-orbiting vehicle is time consuming and costly. It is a labor-intensive, manual process performed pre-mission with the aid of specialized maneuver processing tools. Real-time execution of a rendezvous plan must closely follow a predicted trajectory, and targeted solutions leading up to the terminal phase are computed on the ground. Despite over 25 years of Gemini, Apollo, Skylab, and shuttle vehicle-to-vehicle rendezvous missions flown to date, rendezvous in Earth orbit still requires careful monitoring and cannot be taken for granted. For example, a significant trajectory offset was experienced during terminal phase rendezvous of the STS-32 Long Duration Exposure Facility retrieval mission. Several improvements can be introduced to the present rendezvous planning process to reduce costs, produce more fuel-efficient profiles, and increase the probability of mission success.

GEMINI-TITAN (GT)-10 - EARTH SKY - RENDEZVOUS - OUTER SPACE

NASA Image and Video Library

1966-07-18

S66-46122 (18 July 1966) --- Agena Target Docking Vehicle 5005 is photographed from the Gemini-Titan 10 (GT-10) spacecraft during rendezvous in space. The two spacecraft are about 38 feet apart. After docking with the Agena, astronauts John W. Young, command pilot, and Michael Collins, pilot, fired the 16,000 pound thrust engine of Agena X's primary propulsion system to boost the combined vehicles into an orbit with an apogee of 413 nautical miles to set a new altitude record for manned spaceflight. Photo credit: NASA

On-orbit demonstration of automated closure and capture using ESA-developed proximity operations technologies and an existing, serviceable NASA Explorer Platform spacecraft

NASA Technical Reports Server (NTRS)

Hohwiesner, Bill; Claudinon, Bernard

1991-01-01

The European Space Agency (ESA) has been working to develop an autonomous rendezvous and docking capability since 1984 to enable Hermes to automatically dock with Columbus. As a result, ESA with Matra, MBB, and other space companies have developed technologies that are also directly supportive of the current NASA initiative for Automated Rendezvous and Capture. Fairchild and Matra would like to discuss the results of the applicable ESA/Matra rendezvous and capture developments, and suggest how these capabilities could be used, together with an existing NASA Explorer Platform satellite, to minimize new development and accomplish a cost effective automatic closure and capture demonstration program. Several RV sensors have been developed at breadboard level for the Hermes/Columbus program by Matra, MBB, and SAAB. Detailed algorithms for automatic rendezvous, closure, and capture have been developed by ESA and CNES for application with Hermes to Columbus rendezvous and docking, and they currently are being verified with closed-loop software simulation. The algorithms have multiple closed-loop control modes and phases starting at long range using GPS navigation. Differential navigation is used for coast/continuous thrust homing, holdpoint acquisition, V-bar hopping, and station point acquisition. The proximity operation sensor is used for final closure and capture. A subset of these algorithms, comprising the proximity operations algorithms, could easily be extracted and tailored to a limited objective closure and capture flight demonstration.

Electric Propulsion System Selection Process for Interplanetary Missions

NASA Technical Reports Server (NTRS)

Landau, Damon; Chase, James; Kowalkowski, Theresa; Oh, David; Randolph, Thomas; Sims, Jon; Timmerman, Paul

2008-01-01

The disparate design problems of selecting an electric propulsion system, launch vehicle, and flight time all have a significant impact on the cost and robustness of a mission. The effects of these system choices combine into a single optimization of the total mission cost, where the design constraint is a required spacecraft neutral (non-electric propulsion) mass. Cost-optimal systems are designed for a range of mass margins to examine how the optimal design varies with mass growth. The resulting cost-optimal designs are compared with results generated via mass optimization methods. Additional optimizations with continuous system parameters address the impact on mission cost due to discrete sets of launch vehicle, power, and specific impulse. The examined mission set comprises a near-Earth asteroid sample return, multiple main belt asteroid rendezvous, comet rendezvous, comet sample return, and a mission to Saturn.

Estimating the Mass of Asteroid 433 Eros During the Near Spacecraft Flyby

NASA Technical Reports Server (NTRS)

Yeomans, D.; Antreasian, P.; Cheng, A.; Dunham, D.; Farquhar, R.; Gaskell, R.; Giorgini, J.; Helfrich, C.; Konopliv, A.; McAdams, J.;

Space debris, asteroids and satellite orbits; Proceedings of the Fourth and Thirteenth Workshops, Graz, Austria, June 25-July 7, 1984

NASA Technical Reports Server (NTRS)

Kessler, D. J.; Gruen, E.; Sehnal, L.

1985-01-01

The workshops covered a variety of topics relevant to the identification, characterization and monitoring of near-earth solar system debris. Attention was given to man-made and naturally occurring microparticles, their hazards to present and future spacecraft, and ground- and space-based techniques for tracking both large and small debris. The studies are extended to solid fuel particulates in circular space. Asteroid rendezvous missions are discussed, including propulsion and instrumentation options, the possibility of encountering asteroids during Hohman transfer flights to Venus and/or Mars, and the benefits of multiple encounters by one spacecraft. Finally, equipment and analytical models for generating precise satellite orbits are reviewed.

Mission operations for unmanned nuclear electric propulsion outer planet exploration with a thermionic reactor spacecraft.

NASA Technical Reports Server (NTRS)

Spera, R. J.; Prickett, W. Z.; Garate, J. A.; Firth, W. L.

1971-01-01

Mission operations are presented for comet rendezvous and outer planet exploration NEP spacecraft employing in-core thermionic reactors for electric power generation. The selected reference missions are the Comet Halley rendezvous and a Jupiter orbiter at 5.9 planet radii, the orbit of the moon Io. The characteristics of the baseline multi-mission NEP spacecraft are presented and its performance in other outer planet missions, such as Saturn and Uranus orbiters and a Neptune flyby, are discussed. Candidate mission operations are defined from spacecraft assembly to mission completion. Pre-launch operations are identified. Shuttle launch and subsequent injection to earth escape by the Centaur D-1T are discussed, as well as power plant startup and the heliocentric mission phases. The sequence and type of operations are basically identical for all missions investigated.

Reliable spacecraft rendezvous without velocity measurement

NASA Astrophysics Data System (ADS)

He, Shaoming; Lin, Defu

2018-03-01

This paper investigates the problem of finite-time velocity-free autonomous rendezvous for spacecraft in the presence of external disturbances during the terminal phase. First of all, to address the problem of lack of relative velocity measurement, a robust observer is proposed to estimate the unknown relative velocity information in a finite time. It is shown that the effect of external disturbances on the estimation precision can be suppressed to a relatively low level. With the reconstructed velocity information, a finite-time output feedback control law is then formulated to stabilize the rendezvous system. Theoretical analysis and rigorous proof show that the relative position and its rate can converge to a small compacted region in finite time. Numerical simulations are performed to evaluate the performance of the proposed approach in the presence of external disturbances and actuator faults.

Automated Rendezvous and Docking Sensor Testing at the Flight Robotics Laboratory

NASA Technical Reports Server (NTRS)

Mitchell, J.; Johnston, A.; Howard, R.; Williamson, M.; Brewster, L.; Strack, D.; Cryan, S.

2007-01-01

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as Automated Rendezvous and Docking, AR&D). The crewed versions may also perform AR&D, possibly with a different level of automation and/or autonomy, and must also provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success of the Exploration Program. NASA has the responsibility to determine whether the Crew Exploration Vehicle (CEV) contractor-proposed relative navigation sensor suite will meet the CEV requirements. The relatively low technology readiness of relative navigation sensors for AR&D has been carried as one of the CEV Projects top risks. The AR&D Sensor Technology Project seeks to reduce this risk by increasing technology maturation of selected relative navigation sensor technologies through testing and simulation, and to allow the CEV Project to assess the relative navigation sensors.

Trajectory design for a rendezvous mission to Earth's Trojan asteroid 2010 TK7

NASA Astrophysics Data System (ADS)

Lei, Hanlun; Xu, Bo; Zhang, Lei

2017-12-01

In this paper a rendezvous mission to the Earth's Trojan asteroid 2010 TK7 is proposed, and preliminary transfer trajectories are designed. Due to the high inclination (∼ 20.9°) of the target asteroid relative to the ecliptic plane, direct transfers usually require large amounts of fuel consumption, which is beyond the capacity of current technology. As gravity assist technique could effectively change the inclination of spacecraft's trajectory, it is adopted to reduce the launch energy and rendezvous velocity maneuver. In practical computation, impulsive and low-thrust, gravity-assisted trajectories are considered. Among all the trajectories computed, the low-thrust gravity-assisted trajectory with Venus-Earth-Venus (V-E-V) swingby sequence performs the best in terms of propellant mass. For a spacecraft with initial mass of 800 kg , propellant mass of the best trajectory is 36.74 kg . Numerical results indicate that both the impulsive and low-thrust, gravity-assisted trajectories corresponding to V-E-V sequence could satisfy mission constraints, and can be applied to practical rendezvous mission.

Suboptimal artificial potential function sliding mode control for spacecraft rendezvous with obstacle avoidance

NASA Astrophysics Data System (ADS)

Cao, Lu; Qiao, Dong; Xu, Jingwen

2018-02-01

Sub-Optimal Artificial Potential Function Sliding Mode Control (SOAPF-SMC) is proposed for the guidance and control of spacecraft rendezvous considering the obstacles avoidance, which is derived based on the theories of artificial potential function (APF), sliding mode control (SMC) and state dependent riccati equation (SDRE) technique. This new methodology designs a new improved APF to describe the potential field. It can guarantee the value of potential function converge to zero at the desired state. Moreover, the nonlinear terminal sliding mode is introduced to design the sliding mode surface with the potential gradient of APF, which offer a wide variety of controller design alternatives with fast and finite time convergence. Based on the above design, the optimal control theory (SDRE) is also employed to optimal the shape parameter of APF, in order to add some degree of optimality in reducing energy consumption. The new methodology is applied to spacecraft rendezvous with the obstacles avoidance problem, which is simulated to compare with the traditional artificial potential function sliding mode control (APF-SMC) and SDRE to evaluate the energy consumption and control precision. It is demonstrated that the presented method can avoiding dynamical obstacles whilst satisfying the requirements of autonomous rendezvous. In addition, it can save more energy than the traditional APF-SMC and also have better control accuracy than the SDRE.

A Survey of LIDAR Technology and Its Use in Spacecraft Relative Navigation

NASA Technical Reports Server (NTRS)

Christian, John A.; Cryan, Scott P.

2013-01-01

This paper provides a survey of modern LIght Detection And Ranging (LIDAR) sensors from a perspective of how they can be used for spacecraft relative navigation. In addition to LIDAR technology commonly used in space applications today (e.g. scanning, flash), this paper reviews emerging LIDAR technologies gaining traction in other non-aerospace fields. The discussion will include an overview of sensor operating principles and specific pros/cons for each type of LIDAR. This paper provides a comprehensive review of LIDAR technology as applied specifically to spacecraft relative navigation. HE problem of orbital rendezvous and docking has been a consistent challenge for complex space missions since before the Gemini 8 spacecraft performed the first successful on-orbit docking of two spacecraft in 1966. Over the years, a great deal of effort has been devoted to advancing technology associated with all aspects of the rendezvous, proximity operations, and docking (RPOD) flight phase. After years of perfecting the art of crewed rendezvous with the Gemini, Apollo, and Space Shuttle programs, NASA began investigating the problem of autonomous rendezvous and docking (AR&D) to support a host of different mission applications. Some of these applications include autonomous resupply of the International Space Station (ISS), robotic servicing/refueling of existing orbital assets, and on-orbit assembly.1 The push towards a robust AR&D capability has led to an intensified interest in a number of different sensors capable of providing insight into the relative state of two spacecraft. The present work focuses on exploring the state-of-the-art in one of these sensors - LIght Detection And Ranging (LIDAR) sensors. It should be noted that the military community frequently uses the acronym LADAR (LAser Detection And Ranging) to refer to what this paper calls LIDARs. A LIDAR is an active remote sensing device that is typically used in space applications to obtain the range to one or more points on a target spacecraft. As the name suggests, LIDAR sensors use light (typically a laser) to illuminate the target and measure the time it takes for the emitted signal to return to the sensor. Because the light must travel from the source, to

Apollo-Soyuz test project

NASA Technical Reports Server (NTRS)

1975-01-01

Experiments proposed for the Apollo-Soyuz space mission are discussed. Data focus of space processing and manufacturing, earth surveys, and life sciences. Special efforts were made to test the compatibility of the rendezvous and docking systems for manned spacecraft. Mission planning programs, personnel training, and spacecraft modifications for both spacecraft are included.

Apollo Soyuz test project. USA-USSR, fact sheet

NASA Technical Reports Server (NTRS)

1974-01-01

The Apollo Soyuz Test Project (ASTP) is discussed. The United States and the Soviet Union have agreed to develop compatible rendezvous and docking systems which will provide a basis for docking and rescue on future spacecraft of both nations. The ASTP mission will include testing the rendezvous system in orbit, verifying techniques for transfer of astronauts and cosmonauts, and conducting experiments while docked and undocked. Diagrams of the spacecraft and systems involved in the tests are presented. The prime contractors for the equipment are identified. Biographical data on the astronauts participating in the program are provided.

Rendezvous missions with minimoons from L1

NASA Astrophysics Data System (ADS)

Chyba, M.; Haberkorn, T.; Patterson, G.

2014-07-01

We propose to present asteroid capture missions with the so-called minimoons. Minimoons are small asteroids that are temporarily captured objects on orbits in the Earth-Moon system. It has been suggested that, despite their small capture probability, at any time there are one or two meter diameter minimoons, and progressively greater numbers at smaller diameters. The minimoons orbits differ significantly from elliptical orbits which renders a rendezvous mission more challenging, however they offer many advantages for such missions that overcome this fact. First, they are already on geocentric orbits which results in short duration missions with low Delta-v, this translates in cost efficiency and low-risk targets. Second, beside their close proximity to Earth, an advantage is their small size since it provides us with the luxury to retrieve the entire asteroid and not only a sample of material. Accessing the interior structure of a near-Earth satellite in its morphological context is crucial to an in-depth analysis of the structure of the asteroid. Historically, 2006 RH120 is the only minimoon that has been detected but work is ongoing to determine which modifications to current observation facilities is necessary to provide detection algorithm capabilities. In the event that detection is successful, an efficient algorithm to produce a space mission to rendezvous with the detected minimoon is highly desirable to take advantage of this opportunity. This is the main focus of our work. For the design of the mission we propose the following. The spacecraft is first placed in hibernation on a Lissajoux orbit around the liberation point L1 of the Earth-Moon system. We focus on eight-shaped Lissajoux orbits to take advantage of the stability properties of their invariant manifolds for our transfers since the cost to minimize is the spacecraft fuel consumption. Once a minimoon has been detected we must choose a point on its orbit to rendezvous (in position and velocities) with the spacecraft. This is determined using a combination of distance between the minimoon's orbit to L1 and its energy level with respect to the Lissajoux orbit on which the spacecraft is hibernating. Once the spacecraft rendezvous with the minimoon, it will escort the temporarily captured object to analyze it until the withdrawal time when the spacecraft exits the orbit to return to its hibernating location awaiting for another minimoon to be detected. The entire mission including the return portion can be stated as an optimal control problem, however we choose to break it into smaller sub-problems as a first step to be refined later. To model our control system, we use the circular three-body problem since it provides a good approximation in the vicinity of the Earth-Moon dynamics. Expansion to more refined models will be considered once the problem has been solved for this first approximation. The problem is solved in several steps. First, we consider the time minimal problem since we will use a multiple of it for the minimal fuel consumption problem with fixed time. The techniques used to produce the transfers involve an indirect method based on the necessary optimality condition of the Pontriagyn maximum principle coupled with a continuation method to address the sensitivity of the numerical algorithm to initial values. Time local optimality is verified by computing the Jacobi fields of the Hamiltonian system associated to our optimal control problem to check the second-order conditions of optimality and determine the non-existence of conjugate points.

Genetic algorithm based fuzzy control of spacecraft autonomous rendezvous

NASA Technical Reports Server (NTRS)

Karr, C. L.; Freeman, L. M.; Meredith, D. L.

1990-01-01

The U.S. Bureau of Mines is currently investigating ways to combine the control capabilities of fuzzy logic with the learning capabilities of genetic algorithms. Fuzzy logic allows for the uncertainty inherent in most control problems to be incorporated into conventional expert systems. Although fuzzy logic based expert systems have been used successfully for controlling a number of physical systems, the selection of acceptable fuzzy membership functions has generally been a subjective decision. High performance fuzzy membership functions for a fuzzy logic controller that manipulates a mathematical model simulating the autonomous rendezvous of spacecraft are learned using a genetic algorithm, a search technique based on the mechanics of natural genetics. The membership functions learned by the genetic algorithm provide for a more efficient fuzzy logic controller than membership functions selected by the authors for the rendezvous problem. Thus, genetic algorithms are potentially an effective and structured approach for learning fuzzy membership functions.

Near Earth Asteroid Rendezvous (NEAR) Revised Eros Orbit Phase Trajectory Design

NASA Technical Reports Server (NTRS)

Helfrich, J; Miller, J. K.; Antreasian, P. G.; Carranza, E.; Williams, B. G.; Dunham, D. W.; Farquhar, R. W.; McAdams, J. V.

1999-01-01

Trajectory design of the orbit phase of the NEAR mission involves a new process that departs significantly from those procedures used in previous missions. In most cases, a precise spacecraft ephemeris is designed well in advance of arrival at the target body. For NEAR, the uncertainty in the dynamic environment around Eros does not allow the luxury of a precise spacecraft trajectory to be defined in advance. The principal cause of this uncertainty is the limited knowledge oi' the gravity field a,-id rotational state of Eros. As a result, the concept for the NEAR trajectory design is to define a number of rules for satisfying spacecraft, mission, and science constraints, and then apply these rules to various assumptions for the model of Eros. Nominal, high, and low Eros mass models are used for testing the trajectory design strategy and to bracket the ranges of parameter variations that are expected upon arrival at the asteroid. The final design is completed after arrival at Eros and determination of the actual gravity field and rotational state. As a result of the unplanned termination of the deep space rendezvous maneuver on December 20, 1998, the NEAR spacecraft passed within 3830 km of Eros on December 23, 1998. This flyby provided a brief glimpse of Eros, and allowed for a more accurate model of the rotational parameters and gravity field uncertainty. Furthermore, after the termination of the deep space rendezvous burn, contact with the spacecraft was lost and the NEAR spacecraft lost attitude control. During the subsequent gyrations of the spacecraft, hydrazine thruster firings were used to regain attitude control. This unplanned thruster activity used Much of the fuel margin allocated for the orbit phase. Consequently, minimizing fuel consumption is now even more important.

The comet rendezvous asteroid flyby mission to Comet Kopff - Getting there is half the fun

NASA Technical Reports Server (NTRS)

Sweetser, Theodore H.; Kiedron, Krystyna

1990-01-01

The goal of the Comet Rendezvous Asteroid Flyby mission (CRAF) is to fly 'outward to the beginning', to examine closely what are thought to be remnants of the origins of the solar system. In particular, the CRAF spacecraft will use a two-year delta-V-earth-gravity-assist (delta-V-EGA) trajectory to reach a rendezvous point near the aphelion of the Comet Kopff, flying by the asteroid 449 Hamburga on the way. This paper discusses the trajectory used to get to the comet. Topics covered include the launch period, possible additional asteroid flybys, the earth flyby, the Hamburga flyby, and the rendezvous with Comet Kopff.

Enhanced Gravity Tractor Derived from the Asteroid Redirect Mission for Deflecting Hypothetical Asteroid 2017 PDC

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D.; Reeves, David M.; Abell, Paul A.; Shen, Haijun; Qu, Min

2017-01-01

The Asteroid Redirect Mission (ARM) concept would robotically visit a hazardous-size near-Earth asteroid (NEA) with a rendezvous spacecraft, collect a multi-ton boulder and regolith samples from its surface, demonstrate an innovative planetary defense technique known as the Enhanced Gravity Tractor (EGT), and return the asteroidal material to a stable orbit around the Moon, allowing astronauts to explore the returned material in the mid-2020s. Launch of the robotic vehicle to rendezvous with the ARM reference target, NEA (341843) 2008 EV5, would occur in late 2021 [1,2]. The robotic segment of the ARM concept uses a 40 kW Solar Electric Propulsion (SEP) system with a specific impulse (Isp) of 2600 s, and would provide the first ever demonstration of the EGT technique on a hazardous-size asteroid and validate one method of collecting mass in-situ. The power, propellant, and thrust capability of the ARM robotic spacecraft can be scaled from a 40 kW system to 150 kW and 300 kW, which represent a likely future power level progression. The gravity tractor technique uses the gravitational attraction of a station-keeping spacecraft with the asteroid to provide a velocity change and gradually alter the trajectory of the asteroid. EGT utilizes a spacecraft with a high-efficiency propulsion system, such as Solar Electric Propulsion (SEP), along with mass collected in-situ to augment the mass of the spacecraft, thereby increasing the gravitational force between the objects [3]. As long as the spacecraft has sufficient thrust and propellant capability, the EGT force is only limited by the amount of in-situ mass collected and can be increased several orders of magnitude compared to the traditional gravity tractor technique in which only the spacecraft mass is used to generate the gravitational attraction force. This increase in available force greatly reduces the required deflection time. The collected material can be a single boulder, multiple boulders, regolith, or a combination of different material types using a variety of collection techniques. The EGT concept assumes that the ability to efficiently collect asteroid mass in-situ from a wide variety of asteroid types and environments is a future capability that will be developed and perfected in the future by the asteroid mining community. Additionally, it is anticipated that the mass collection would likely be performed by a single or multiple separable spacecraft to allow the SEP spacecraft to operate at safe distance from the asteroid.

Holographic Weapons Sight as Crew Optical Alignment Sight

NASA Technical Reports Server (NTRS)

Merancy, Nujoud; Dehmlow, Brian; Brazzel, Jack P.

2011-01-01

Crew Optical Alignment Sights (COAS) are used by spacecraft pilots to provide a visual reference to a target spacecraft for lateral relative position during rendezvous and docking operations. NASA s Orion vehicle, which is currently under development, has not included a COAS in favor of automated sensors, but the crew office has requested such a device be added for situational awareness and contingency support. The current Space Shuttle COAS was adopted from Apollo heritage, weighs several pounds, and is no longer available for procurement which would make re-use difficult. In response, a study was conducted to examine the possibility of converting a commercially available weapons sight to a COAS for the Orion spacecraft. The device used in this study was the XPS series Holographic Weapon Sight (HWS) procured from L-3 EOTech. This device was selected because the targeting reticule can subtend several degrees, and display a graphic pattern tailored to rendezvous and docking operations. Evaluations of the COAS were performed in both the Orion low-fidelity mockup and rendezvous simulations in the Reconfigurable Operational Cockpit (ROC) by crewmembers, rendezvous engineering experts, and flight controllers at Johnson Space Center. These evaluations determined that this unit s size and mounting options can support proper operation and that the reticule visual qualities are as good as or better than the current Space Shuttle COAS. The results positively indicate that the device could be used as a functional COAS and supports a low-cost technology conversion solution.

Imaging Flash Lidar for Safe Landing on Solar System Bodies and Spacecraft Rendezvous and Docking

NASA Technical Reports Server (NTRS)

Amzajerdian, Farzin; Roback, Vincent E.; Bulyshev, Alexander E.; Brewster, Paul F.; Carrion, William A; Pierrottet, Diego F.; Hines, Glenn D.; Petway, Larry B.; Barnes, Bruce W.; Noe, Anna M.

2015-01-01

NASA has been pursuing flash lidar technology for autonomous, safe landing on solar system bodies and for automated rendezvous and docking. During the final stages of the landing from about 1 kilometer to 500 meters above the ground, the flash lidar can generate 3-Dimensional images of the terrain to identify hazardous features such as craters, rocks, and steep slopes. The onboard flight computer can then use the 3-D map of terrain to guide the vehicle to a safe location. As an automated rendezvous and docking sensor, the flash lidar can provide relative range, velocity, and bearing from an approaching spacecraft to another spacecraft or a space station. NASA Langley Research Center has developed and demonstrated a flash lidar sensor system capable of generating 16,000 pixels range images with 7 centimeters precision, at 20 Hertz frame rate, from a maximum slant range of 1800 m from the target area. This paper describes the lidar instrument and presents the results of recent flight tests onboard a rocket-propelled free-flyer vehicle (Morpheus) built by NASA Johnson Space Center. The flights were conducted at a simulated lunar terrain site, consisting of realistic hazard features and designated landing areas, built at NASA Kennedy Space Center specifically for this demonstration test. This paper also provides an overview of the plan for continued advancement of the flash lidar technology aimed at enhancing its performance to meet both landing and automated rendezvous and docking applications.

Optimised collision avoidance for an ultra-close rendezvous with a failed satellite based on the Gauss pseudospectral method

NASA Astrophysics Data System (ADS)

Chu, Xiaoyu; Zhang, Jingrui; Lu, Shan; Zhang, Yao; Sun, Yue

2016-11-01

This paper presents a trajectory planning algorithm to optimise the collision avoidance of a chasing spacecraft operating in an ultra-close proximity to a failed satellite. The complex configuration and the tumbling motion of the failed satellite are considered. The two-spacecraft rendezvous dynamics are formulated based on the target body frame, and the collision avoidance constraints are detailed, particularly concerning the uncertainties. An optimisation solution of the approaching problem is generated using the Gauss pseudospectral method. A closed-loop control is used to track the optimised trajectory. Numerical results are provided to demonstrate the effectiveness of the proposed algorithms.

Robust control for spacecraft rendezvous system with actuator unsymmetrical saturation: a gain scheduling approach

NASA Astrophysics Data System (ADS)

Wang, Qian; Xue, Anke

2018-06-01

This paper has proposed a robust control for the spacecraft rendezvous system by considering the parameter uncertainties and actuator unsymmetrical saturation based on the discrete gain scheduling approach. By changing of variables, we transform the actuator unsymmetrical saturation control problem into a symmetrical one. The main advantage of the proposed method is improving the dynamic performance of the closed-loop system with a region of attraction as large as possible. By the Lyapunov approach and the scheduling technology, the existence conditions for the admissible controller are formulated in the form of linear matrix inequalities. The numerical simulation illustrates the effectiveness of the proposed method.

Spacecraft rendezvous operational considerations affecting vehicle systems design and configuration

NASA Astrophysics Data System (ADS)

Prust, Ellen E.

One lesson learned from Orbiting Maneuvering Vehicle (OMV) program experience is that Design Reference Missions must include an appropriate balance of operations and performance inputs to effectively drive vehicle systems design and configuration. Rendezvous trajectory design is based on vehicle characteristics (e.g., mass, propellant tank size, and mission duration capability) and operational requirements, which have evolved through the Gemini, Apollo, and STS programs. Operational constraints affecting the rendezvous final approach are summarized. The two major objectives of operational rendezvous design are vehicle/crew safety and mission success. Operational requirements on the final approach which support these objectives include: tracking/targeting/communications; trajectory dispersion and navigation uncertainty handling; contingency protection; favorable sunlight conditions; acceptable relative state for proximity operations handover; and compliance with target vehicle constraints. A discussion of the ways each of these requirements may constrain the rendezvous trajectory follows. Although the constraints discussed apply to all rendezvous, the trajectory presented in 'Cargo Transfer Vehicle Preliminary Reference Definition' (MSFC, May 1991) was used as the basis for the comments below.

Optimal cooperative time-fixed impulsive rendezvous

NASA Technical Reports Server (NTRS)

Mirfakhraie, Koorosh; Conway, Bruce A.; Prussing, John E.

1988-01-01

A method has been developed for determining optimal, i.e., minimum fuel, trajectories for the fixed-time cooperative rendezvous of two spacecraft. The method presently assumes that the vehicles perform a total of three impulsive maneuvers with each vehicle being active, that is, making at least one maneuver. The cost of a feasible 'reference' trajectory is improved by an optimizer which uses an analytical gradient developed using primer vector theory and a new solution for the optimal terminal (rendezvous) maneuver. Results are presented for a large number of cases in which the initial orbits of both vehicles are circular but in which the initial positions of the vehicles and the allotted time for rendezvous are varied. In general, the cost of the cooperative rendezvous is less than that of rendezvous with one vehicle passive. Further improvement in cost may be obtained in the future when additional, i.e., midcourse, impulses are allowed and inserted as indicated for some cases by the primer vector histories which are generated by the program.

Preflight SL-1/SL-3 Skylab VHF ranging coverage (nominal TPI). Antenna and propagation studies for spacecraft systems, task E-531

NASA Technical Reports Server (NTRS)

Eisenhauer, D. R.; James, D. A.

1973-01-01

A preflight assessment of the Skylab VHF ranging coverage for the rendezvous portion of the nominal SL-1/SL-3 mission is reported, assuming a 27 July 1973 SL-3 launch. Data are based on a nominal attitude trajectory, which has the Saturn workshop in a solar inertial attitude throughout the rendezvous; the CSM terminal phase initiation maneuver is nominal. An addendum to this report is being prepared, which considers the effects of early and late TPI maneuvers. Curves are presented which show the variation in received power levels on both spacecraft-to-spacecraft links from about 600 n.mi. range to CSM and SWS station keeping. Appropriate threshold levels are shown on these received power curves to indicate zero circuit margins for the ranging function.

Demonstration of Autonomous Rendezvous Technology (DART) Project Summary

NASA Technical Reports Server (NTRS)

Rumford, TImothy E.

2003-01-01

Since the 1960's, NASA has performed numerous rendezvous and docking missions. The common element of all US rendezvous and docking is that the spacecraft has always been piloted by astronauts. Only the Russian Space Program has developed and demonstrated an autonomous capability. The Demonstration of Autonomous Rendezvous Technology (DART) project currently funded under NASA's Space Launch Initiative (SLI) Cycle I, provides a key step in establishing an autonomous rendezvous capability for the United States. DART's objective is to demonstrate, in space, the hardware and software necessary for autonomous rendezvous. Orbital Sciences Corporation intends to integrate an Advanced Video Guidance Sensor and Autonomous Rendezvous and Proximity Operations algorithms into a Pegasus upper stage in order to demonstrate the capability to autonomously rendezvous with a target currently in orbit. The DART mission will occur in April 2004. The launch site will be Vandenburg AFB and the launch vehicle will be a Pegasus XL equipped with a Hydrazine Auxiliary Propulsion System 4th stage. All mission objectives will be completed within a 24 hour period. The paper provides a summary of mission objectives, mission overview and a discussion on the design features of the chase and target vehicles.

Development of a cooperative operational rendezvous plan for Eureca and other maneuvering Shuttle payloads

NASA Technical Reports Server (NTRS)

Gavin, R. T.

1987-01-01

This paper discusses the development of a new class of US Space Shuttle rendezvous missions which involve a maneuvering target vehicle. The objective of the analysis was to develop an operational plan to take advantage of the target spacecraft's maneuvering ability by making it responsible for a portion of the maneuvers necessary to achieve rendezvous. This work resulted in the development of a region in space relative to the Shuttle, called the control box, into which the target vehicle maneuvers. Furthermore, a mission operations plan was developed to implement the control box technique.

Development of an autonomous video rendezvous and docking system, phase 2

NASA Technical Reports Server (NTRS)

Tietz, J. C.; Richardson, T. E.

1983-01-01

The critical elements of an autonomous video rendezvous and docking system were built and used successfully in a physical laboratory simulation. The laboratory system demonstrated that a small, inexpensive electronic package and a flight computer of modest size can analyze television images to derive guidance information for spacecraft. In the ultimate application, the system would use a docking aid consisting of three flashing lights mounted on a passive target spacecraft. Television imagery of the docking aid would be processed aboard an active chase vehicle to derive relative positions and attitudes of the two spacecraft. The demonstration system used scale models of the target spacecraft with working docking aids. A television camera mounted on a 6 degree of freedom (DOF) simulator provided imagery of the target to simulate observations from the chase vehicle. A hardware video processor extracted statistics from the imagery, from which a computer quickly computed position and attitude. Computer software known as a Kalman filter derived velocity information from position measurements.

The asteroid rendezvous spacecraft. An adaptation study of TIROS/DMSP technology

NASA Technical Reports Server (NTRS)

1982-01-01

The feasibility of using the TIROS/DMSP Earth orbiting meteorological satellite in application to a near Earth asteroid rendezvous mission. System and subsystems analysis was carried out to develop a configuration of the spacecraft suitable for this mission. Mission analysis studies were also done and maneuver/rendezvous scenarios developed for baseline missions to both Anteros and Eros. The fact that the Asteroid mission is the most complex of the Pioneer class missions currently under consideration notwithstanding, the basic conclusion very strongly supports the suitability of the basic TIROS bus for this mission in all systems and subsystems areas, including science accommodation. Further, the modifications which are required due to the unique mission are very low risk and can be accomplished readily. The key issue is that in virtually every key subsystem, the demands of the Asteroid mission are a subset of the basic meteorological satellite mission. This allows a relatively simple reconfiguration to be accomplished without a major system redesign.

Solar array study for solar electric propulsion spacecraft for the Encke rendezvous mission

NASA Technical Reports Server (NTRS)

Sequeira, E. A.; Patterson, R. E.

1974-01-01

The work is described which was performed on the design, analysis and performance of a 20 kW rollup solar array capable of meeting the design requirements of a solar electric spacecraft for the 1980 Encke rendezvous mission. To meet the high power requirements of the proposed electric propulsion mission, solar arrays on the order of 186.6 sq m were defined. Because of the large weights involved with arrays of this size, consideration of array configurations is limited to lightweight, large area concepts with maximum power-to-weight ratios. Items covered include solar array requirements and constraints, array concept selection and rationale, structural and electrical design considerations, and reliability considerations.

Asteroid Impact Mission (aim) & Deflection Assessment: AN Opportunity to Understand Impact Dynamics and Modelling

NASA Astrophysics Data System (ADS)

Galvez, A.; Carnelli, I.; Fontaine, M.; Corral Van Damme, C.

2012-09-01

ESA's Future Preparation and Strategic Studies Office has carried out the Asteroid Impact Mission (AIM) study with the objective of defining an affordable and fully independent mission element that ESA could contribute to an Asteroid Impact Deflection Assessment campaign (AIDA), a joint effort of ESA, JHU/APL, NASA, OCA and DLR. The mission design foresees two independent spacecraft, one impactor (DART) and one rendezvous probe (AIM). The target of this mission is the binary asteroid system (65803) Didymos (1996 GT): one spacecraft, DART, would impact the secondary of the Didymos binary system while AIM would observe and measure any the change in the relative orbit. For this joint project, the timing of the experiment is set (maximum proximity of the target to Earth allowing for ground-based characterisation of the experiment) but the spacecraft are still able to pursue their missions fully independently. This paper describes in particular the AIM rendezvous mission concept.

Orbital Express fluid transfer demonstration system

NASA Astrophysics Data System (ADS)

Rotenberger, Scott; SooHoo, David; Abraham, Gabriel

2008-04-01

Propellant resupply of orbiting spacecraft is no longer in the realm of high risk development. The recently concluded Orbital Express (OE) mission included a fluid transfer demonstration that operated the hardware and control logic in space, bringing the Technology Readiness Level to a solid TRL 7 (demonstration of a system prototype in an operational environment). Orbital Express (funded by the Defense Advanced Research Projects Agency, DARPA) was launched aboard an Atlas-V rocket on March 9th, 2007. The mission had the objective of demonstrating technologies needed for routine servicing of spacecraft, namely autonomous rendezvous and docking, propellant resupply, and orbital replacement unit transfer. The demonstration system used two spacecraft. A servicing vehicle (ASTRO) performed multiple dockings with the client (NextSat) spacecraft, and performed a variety of propellant transfers in addition to exchanges of a battery and computer. The fluid transfer and propulsion system onboard ASTRO, in addition to providing the six degree-of-freedom (6 DOF) thruster system for rendezvous and docking, demonstrated autonomous transfer of monopropellant hydrazine to or from the NextSat spacecraft 15 times while on orbit. The fluid transfer system aboard the NextSat vehicle was designed to simulate a variety of client systems, including both blowdown pressurization and pressure regulated propulsion systems. The fluid transfer demonstrations started with a low level of autonomy, where ground controllers were allowed to review the status of the demonstration at numerous points before authorizing the next steps to be performed. The final transfers were performed at a full autonomy level where the ground authorized the start of a transfer sequence and then monitored data as the transfer proceeded. The major steps of a fluid transfer included the following: mate of the coupling, leak check of the coupling, venting of the coupling, priming of the coupling, fluid transfer, gauging of receiving tank, purging of coupling and de-mate of the coupling.

Spacecraft Mission Design for the Mitigation of the 2017 PDC Hypothetical Asteroid Threat

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Sarli, Bruno V.; Lyzhoft, Josh; Chodas, Paul W.; Englander, Jacob A.

2017-01-01

This paper presents a detailed mission design analysis results for the 2017 Planetary Defense Conference (PDC) Hypothetical Asteroid Impact Scenario, documented at https:cneos.jpl.nasa.govpdcspdc17. The mission design includes campaigns for both reconnaissance (flyby or rendezvous) of the asteroid (to characterize it and the nature of the threat it poses to Earth) and mitigation of the asteroid, via kinetic impactor deflection, nuclear explosive device (NED) deflection, or NED disruption. Relevant scenario parameters are varied to assess the sensitivity of the design outcome, such as asteroid bulk density, asteroid diameter, momentum enhancement factor, spacecraft launch vehicle, and mitigation system type. Different trajectory types are evaluated in the mission design process from purely ballistic to those involving optimal midcourse maneuvers, planetary gravity assists, and/or low-thrust solar electric propulsion. The trajectory optimization is targeted around peak deflection points that were found through a novel linear numerical technique method. The optimization process includes constrain parameters, such as Earth departure date, launch declination, spacecraft, asteroid relative velocity and solar phase angle, spacecraft dry mass, minimum/maximum spacecraft distances from Sun and Earth, and Earth-spacecraft communications line of sight. Results show that one of the best options for the 2017 PDC deflection is solar electric propelled rendezvous mission with a single spacecraft using NED for the deflection.

CoMA: A high resolution Time-Of-Flight Secondary Ion Mass Spectrometer (TOF-SIMS) for in situ analysis of cometary matter

NASA Technical Reports Server (NTRS)

Zscheeg, Harry; Kissel, J.; Natour, G.

1992-01-01

A lot of clues concerning the origin of the solar system can be found by sending an exploring spacecraft to a rendezvous with a comet. The space experiment CoMA, which will measure the elemental, isotopic, and molecular composition of cometary dust grains is described. It will be flown on NASA's Comet Rendezvous Asteroid Flyby (CRAF) mission.

Orion Handling Qualities During ISS Rendezvous and Docking

NASA Technical Reports Server (NTRS)

Hart, Jeremy J.; Stephens, J. P.; Spehar, P.; Bilimoria, K.; Foster, C.; Gonzalex, R.; Sullivan, K.; Jackson, B.; Brazzel, J.; Hart, J.

2011-01-01

The Orion spacecraft was designed to rendezvous with multiple vehicles in low earth orbit (LEO) and beyond. To perform the required rendezvous and docking task, Orion must provide enough control authority to perform coarse translational maneuvers while maintaining precision to perform the delicate docking corrections. While Orion has autonomous docking capabilities, it is expected that final approach and docking operations with the International Space Station (ISS) will initially be performed in a manual mode. A series of evaluations was conducted by NASA and Lockheed Martin at the Johnson Space Center to determine the handling qualities (HQ) of the Orion spacecraft during different docking and rendezvous conditions using the Cooper-Harper scale. This paper will address the specifics of the handling qualities methodology, vehicle configuration, scenarios flown, data collection tools, and subject ratings and comments. The initial Orion HQ assessment examined Orion docking to the ISS. This scenario demonstrates the Translational Hand Controller (THC) handling qualities of Orion. During this initial assessment, two different scenarios were evaluated. The first was a nominal docking approach to a stable ISS, with Orion initializing with relative position dispersions and a closing rate of approximately 0.1 ft/sec. The second docking scenario was identical to the first, except the attitude motion of the ISS was modeled to simulate a stress case ( 1 degree deadband per axis and 0.01 deg/sec rate deadband per axis). For both scenarios, subjects started each run on final approach at a docking port-to-port range of 20 ft. Subjects used the THC in pulse mode with cues from the docking camera image, window views, and range and range rate data displayed on the Orion display units. As in the actual design, the attitude of the Orion vehicle was held by the automated flight control system at 0.5 degree deadband per axis. Several error sources were modeled including Reaction Control System (RCS) jet angular and position misalignment, RCS thrust magnitude uncertainty, RCS jet force direction uncertainty due to self plume impingement, and Orion center of mass uncertainty.

Concept designs for NASA's Solar Electric Propulsion Technology Demonstration Mission

NASA Technical Reports Server (NTRS)

Mcguire, Melissa L.; Hack, Kurt J.; Manzella, David H.; Herman, Daniel A.

2014-01-01

Multiple Solar Electric Propulsion Technology Demonstration Mission were developed to assess vehicle performance and estimated mission cost. Concepts ranged from a 10,000 kilogram spacecraft capable of delivering 4000 kilogram of payload to one of the Earth Moon Lagrange points in support of future human-crewed outposts to a 180 kilogram spacecraft capable of performing an asteroid rendezvous mission after launched to a geostationary transfer orbit as a secondary payload. Low-cost and maximum Delta-V capability variants of a spacecraft concept based on utilizing a secondary payload adapter as the primary bus structure were developed as were concepts designed to be co-manifested with another spacecraft on a single launch vehicle. Each of the Solar Electric Propulsion Technology Demonstration Mission concepts developed included an estimated spacecraft cost. These data suggest estimated spacecraft costs of $200 million - $300 million if 30 kilowatt-class solar arrays and the corresponding electric propulsion system currently under development are used as the basis for sizing the mission concept regardless of launch vehicle costs. The most affordable mission concept developed based on subscale variants of the advanced solar arrays and electric propulsion technology currently under development by the NASA Space Technology Mission Directorate has an estimated cost of $50M and could provide a Delta-V capability comparable to much larger spacecraft concepts.

Terminal spacecraft rendezvous and capture with LASSO model predictive control

NASA Astrophysics Data System (ADS)

Hartley, Edward N.; Gallieri, Marco; Maciejowski, Jan M.

2013-11-01

The recently investigated ℓasso model predictive control (MPC) is applied to the terminal phase of a spacecraft rendezvous and capture mission. The interaction between the cost function and the treatment of minimum impulse bit is also investigated. The propellant consumption with ℓasso MPC for the considered scenario is noticeably less than with a conventional quadratic cost and control actions are sparser in time. Propellant consumption and sparsity are competitive with those achieved using a zone-based ℓ1 cost function, whilst requiring fewer decision variables in the optimisation problem than the latter. The ℓasso MPC is demonstrated to meet tighter specifications on control precision and also avoids the risk of undesirable behaviours often associated with pure ℓ1 stage costs.

Study of a comet rendezvous mission, volume 1

NASA Technical Reports Server (NTRS)

1972-01-01

The feasibility, scientific objectives, modes of exploration and implementation alternatives of a rendezvous mission to Encke's comet in 1984 are considered. Principal emphasis is placed on developing the scientific rationale for such a mission, based on available knowledge and best estimates of this comet's physical characteristics, including current theories of its origin, evolution and composition. Studied are mission profile alternatives, performance tradeoffs, preferred exploration strategy, and a spacecraft design concept capable of performing this mission. The study showed that the major scientific objectives can be met by a Titan IIID/Centaur-launched 17.5 kw solar electric propulsion spacecraft which carries 60 kg of scientific instruments and is capable of extensive maneuvering within the comet envelope to explore the coma, tail and nucleus.

Polarized-interferometer feasibility study

NASA Technical Reports Server (NTRS)

Raab, F. H.

1983-01-01

The feasibility of using a polarized-interferometer system as a rendezvous and docking sensor for two cooperating spacecraft was studied. The polarized interferometer is a radio frequency system for long range, real time determination of relative position and attitude. Range is determined by round trip signal timing. Direction is determined by radio interferometry. Relative roll is determined from signal polarization. Each spacecraft is equipped with a transponder and an antenna array. The antenna arrays consist of four crossed dipoles that can transmit or receive either circularly or linearly polarized signals. The active spacecraft is equipped with a sophisticated transponder and makes all measurements. The transponder on the passive spacecraft is a relatively simple repeater. An initialization algorithm is developed to estimate position and attitude without any a priori information. A tracking algorithm based upon minimum variance linear estimators is also developed. Techniques to simplify the transponder on the passive spacecraft are investigated and a suitable configuration is determined. A multiple carrier CW signal format is selected. The dependence of range accuracy and ambiguity resolution error probability are derived and used to design a candidate system. The validity of the design and the feasibility of the polarized interferometer concept are verified by simulation.

Concurrent engineering: Spacecraft and mission operations system design

NASA Technical Reports Server (NTRS)

Landshof, J. A.; Harvey, R. J.; Marshall, M. H.

1994-01-01

Despite our awareness of the mission design process, spacecraft historically have been designed and developed by one team and then turned over as a system to the Mission Operations organization to operate on-orbit. By applying concurrent engineering techniques and envisioning operability as an essential characteristic of spacecraft design, tradeoffs can be made in the overall mission design to minimize mission lifetime cost. Lessons learned from previous spacecraft missions will be described, as well as the implementation of concurrent mission operations and spacecraft engineering for the Near Earth Asteroid Rendezvous (NEAR) program.

Hardware-in-the-Loop Rendezvous Tests of a Novel Actuators Command Concept

NASA Astrophysics Data System (ADS)

Gomes dos Santos, Willer; Marconi Rocco, Evandro; Boge, Toralf; Benninghoff, Heike; Rems, Florian

2016-12-01

Integration, test and validation results, in a real-time environment, of a novel concept for spacecraft control are presented in this paper. The proposed method commands simultaneously a group of actuators optimizing a given set of objective functions based on a multiobjective optimization technique. Since close proximity maneuvers play an important role in orbital servicing missions, the entire GNC system has been integrated and tested at a hardware-in-the-loop (HIL) rendezvous and docking simulator known as European Proximity Operations Simulator (EPOS). During the test campaign at EPOS facility, a visual camera has been used to provide the necessary measurements for calculating the relative position with respect to the target satellite during closed-loop simulations. In addition, two different configurations of spacecraft control have been considered in this paper: a thruster reaction control system and a mixed actuators mode which includes thrusters, reaction wheels, and magnetic torqrods. At EPOS, results of HIL closed-loop tests have demonstrated that a safe and stable rendezvous approach can be achieved with the proposed GNC loop.

Spacecraft Mission Design for the Mitigation of the 2017 PDC Hypothetical Asteroid Threat

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Sarli, Bruno V.; Lyzhoft, Joshua; Chodas, Paul W.; Englander, Jacob A.

2017-01-01

This paper presents a detailed mission design analysis results for the 2017 Planetary Defense Conference (PDC) Hypothetical Asteroid Impact Scenario, documented at https://cneos.jpl.nasa.gov/ pd/cs/pdc17/. The mission design includes campaigns for both reconnaissance (flyby or rendezvous) of the asteroid (to characterize it and the nature of the threat it poses to Earth) and mitigation of the asteroid, via kinetic impactor deflection, nuclear explosive device (NED) deflection, or NED disruption. Relevant scenario parameters are varied to assess the sensitivity of the design outcome, such as asteroid bulk density, asteroid diameter, momentum enhancement factor, spacecraft launch vehicle, and mitigation system type. Different trajectory types are evaluated in the mission design process from purely ballistic to those involving optimal midcourse maneuvers, planetary gravity assists, and/or lowthrust solar electric propulsion. The trajectory optimization is targeted around peak deflection points that were found through a novel linear numerical technique method. The optimization process includes constrain parameters, such as Earth departure date, launch declination, spacecraft/asteroid relative velocity and solar phase angle, spacecraft dry mass, minimum/maximum spacecraft distances from Sun and Earth, and Earth/spacecraft communications line of sight. Results show that one of the best options for the 2017 PDC deflection is solar electric propelled rendezvous mission with a single spacecraft using NED for the deflection

Comet rendezvous mission design using Solar Electric Propulsion

NASA Technical Reports Server (NTRS)

Sackett, L. L.; Hastrup, R. C.; Yen, C.-W. L.; Wood, L. J.

1979-01-01

A dual comet (Halley Flyby/Tempel 2 Rendezvous) mission, which is planned to be the first to use the Solar Electric Propulsion System (SEPS), is to be launched in 1985. The purpose of this paper is to describe how the mission design attempts to maximize science return while working within spacecraft and other constraints. Science requirements and desires are outlined and specific instruments are considered. Emphasis is on the strategy for operations in the vicinity of Tempel 2, for which a representative profile is described. The mission is planned to extend about one year past initial rendezvous. Because of the large uncertainty in the comet environment the Tempel 2 operations strategy must be highly adaptive.

Progress in navigation filter estimate fusion and its application to spacecraft rendezvous

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell

1994-01-01

A new derivation of an algorithm which fuses the outputs of two Kalman filters is presented within the context of previous research in this field. Unlike other works, this derivation clearly shows the combination of estimates to be optimal, minimizing the trace of the fused covariance matrix. The algorithm assumes that the filters use identical models, and are stable and operating optimally with respect to their own local measurements. Evidence is presented which indicates that the error ellipsoid derived from the covariance of the optimally fused estimate is contained within the intersections of the error ellipsoids of the two filters being fused. Modifications which reduce the algorithm's data transmission requirements are also presented, including a scalar gain approximation, a cross-covariance update formula which employs only the two contributing filters' autocovariances, and a form of the algorithm which can be used to reinitialize the two Kalman filters. A sufficient condition for using the optimally fused estimates to periodically reinitialize the Kalman filters in this fashion is presented and proved as a theorem. When these results are applied to an optimal spacecraft rendezvous problem, simulated performance results indicate that the use of optimally fused data leads to significantly improved robustness to initial target vehicle state errors. The following applications of estimate fusion methods to spacecraft rendezvous are also described: state vector differencing, and redundancy management.

Optimal four-impulse rendezvous between coplanar elliptical orbits

NASA Astrophysics Data System (ADS)

Wang, JianXia; Baoyin, HeXi; Li, JunFeng; Sun, FuChun

2011-04-01

Rendezvous in circular or near circular orbits has been investigated in great detail, while rendezvous in arbitrary eccentricity elliptical orbits is not sufficiently explored. Among the various optimization methods proposed for fuel optimal orbital rendezvous, Lawden's primer vector theory is favored by many researchers with its clear physical concept and simplicity in solution. Prussing has applied the primer vector optimization theory to minimum-fuel, multiple-impulse, time-fixed orbital rendezvous in a near circular orbit and achieved great success. Extending Prussing's work, this paper will employ the primer vector theory to study trajectory optimization problems of arbitrary eccentricity elliptical orbit rendezvous. Based on linearized equations of relative motion on elliptical reference orbit (referred to as T-H equations), the primer vector theory is used to deal with time-fixed multiple-impulse optimal rendezvous between two coplanar, coaxial elliptical orbits with arbitrary large eccentricity. A parameter adjustment method is developed for the prime vector to satisfy the Lawden's necessary condition for the optimal solution. Finally, the optimal multiple-impulse rendezvous solution including the time, direction and magnitudes of the impulse is obtained by solving the two-point boundary value problem. The rendezvous error of the linearized equation is also analyzed. The simulation results confirmed the analyzed results that the rendezvous error is small for the small eccentricity case and is large for the higher eccentricity. For better rendezvous accuracy of high eccentricity orbits, a combined method of multiplier penalty function with the simplex search method is used for local optimization. The simplex search method is sensitive to the initial values of optimization variables, but the simulation results show that initial values with the primer vector theory, and the local optimization algorithm can improve the rendezvous accuracy effectively with fast convergence, because the optimal results obtained by the primer vector theory are already very close to the actual optimal solution. If the initial values are taken randomly, it is difficult to converge to the optimal solution.

A Reconfigurable Testbed Environment for Spacecraft Autonomy

NASA Technical Reports Server (NTRS)

Biesiadecki, Jeffrey; Jain, Abhinandan

1996-01-01

A key goal of NASA's New Millennium Program is the development of technology for increased spacecraft on-board autonomy. Achievement of this objective requires the development of a new class of ground-based automony testbeds that can enable the low-cost and rapid design, test, and integration of the spacecraft autonomy software. This paper describes the development of an Autonomy Testbed Environment (ATBE) for the NMP Deep Space I comet/asteroid rendezvous mission.

GEMINI-9 - EARTH SKY - ATDA

NASA Image and Video Library

1966-06-06

S66-37972 (3 June 1966) ?-- The Augmented Target Docking Adapter (ATDA) is photographed from the Gemini-9 spacecraft during one of three rendezvous occasions in space. The ATDA and Gemini-9 spacecraft are 35.5 feet apart in this view. Failure of the docking adapter protective cover on the ATDA to fully separate prevented the docking of the two spacecraft. The ATDA was described by the Gemini-9 crew members as an ?angry alligator.? Photo credit: NASA

A Concept for In-space, System-level Validation of Spacecraft Precision Formation Flying

NASA Technical Reports Server (NTRS)

Leitner, Jesse; Carpenter, J. Russell; Naasz, Bo J.; Scharf, Daniel P.; Hadaegh, Fred Y.; Ahmed, Asif

2007-01-01

A number of international space agencies and organizations, to include the National Aeronautics and Space Administration (NASA), the European Space Agency (ESA), and the Centre National d'Etudes Spatiales (CNES), to name a few, have embraced the concept of spacecraft formation flying to revolutionize the capabilities of astronomy and Earth remote sensing from space. The concept has been around well over a decade and a wide array of technologies and capabilities have been developed to enable multiple spacecraft to collaborate in a highly-coupled manner as would be required for a formation flying mission. Furthermore, many relevant capabilities for formation flying have been demonstrated in the area of rendezvous and docking, loosely-controlled formations, and in missions with collaborating spacecraft with very precise metrology. .However, in considering the case of precision formation flying (PFF), i.e, when the relative geometry of multiple vehicles must be controlled on-board in a continuous and precise manner, there have been several missions proposed, but the realization in space has not yet occurred due to a range of issues. This paper will briefly examine those issues and present a concept for demonstrating a core capability for performing PFF, necessary for virtually any PFF mission concept, that will help to overcome the problems encountered in prior attempts and help to allay the risks to enable future PFF science missions.

Spacecraft formation flying for Earth-crossing object deflections using a power limited laser ablating

NASA Astrophysics Data System (ADS)

Yoo, Sung-Moon; Song, Young-Joo; Park, Sang-Young; Choi, Kyu-Hong

2009-06-01

A formation flying strategy with an Earth-crossing object (ECO) is proposed to avoid the Earth collision. Assuming that a future conceptual spacecraft equipped with a powerful laser ablation tool already rendezvoused with a fictitious Earth collision object, the optimal required laser operating duration and direction histories are accurately derived to miss the Earth. Based on these results, the concept of formation flying between the object and the spacecraft is applied and analyzed as to establish the spacecraft's orbital motion design strategy. A fictitious "Apophis"-like object is established to impact with the Earth and two major deflection scenarios are designed and analyzed. These scenarios include the cases for the both short and long laser operating duration to avoid the Earth impact. Also, requirement of onboard laser tool's for both cases are discussed. As a result, the optimal initial conditions for the spacecraft to maintain its relative trajectory to the object are discovered. Additionally, the discovered optimal initial conditions also satisfied the optimal required laser operating conditions with no additional spacecraft's own fuel expenditure to achieve the spacecraft formation flying with the ECO. The initial conditions founded in the current research can be used as a spacecraft's initial rendezvous points with the ECO when designing the future deflection missions with laser ablation tools. The results with proposed strategy are expected to make more advances in the fields of the conceptual studies, especially for the future deflection missions using powerful laser ablation tools.

Phobos/Deimos sample return via solar sail.

PubMed

Matloff, Gregory L; Taylor, Travis; Powell, Conley; Moton, Tryshanda

2005-12-01

A sample-return mission to the Martian satellites using a con-temporary solar sail for all post-Earth-escape propulsion is proposed. The 0.015 kg/m(2) areal mass-thickness sail unfurls after launch and injection onto a Mars-bound Hohmann-transfer ellipse. Structure and payload increase spacecraft areal mass thickness to 0.028 kg/m(2). During the Mars encounter, the sail functions as a parachute in the outer atmosphere of Mars to accomplish aerocapture. On-board thrusters or the sail maneuver the spacecraft into an orbit with periapsis near Mars and apoapsis near Phobos. The orbit is circularized for Phobos-rendezvous; surface samples are collected. The sail then raises the orbit for Deimos-rendezvous and sample collection. The sail next places the spacecraft on an Earth-bound Hohmann-transfer ellipse. During Earth encounter, the sail accomplishes Earth-aerocapture or partially decelerates the sample container for entry into the Earth's atmosphere. Mission mass budget is about 218 grams and mission duration is less than five years.

Phobos/Deimos Sample Return via Solar Sail

NASA Technical Reports Server (NTRS)

Matloff, Gregory L.; Taylor, Travis; Powell, Conley; Moton, Tryshanda

2004-01-01

Abstract A sample-return mission to the martian satellites using a contemporary solar sail for all post-Earth-escape propulsion is proposed. The 0.015 kg/sq m areal mass-thickness sail unfurls after launch and injection onto a Mars-bound Hohmann-transfer ellipse. Structure and pay!oad increase spacecraft areal mass thickness to 0.028 kg/sq m. During Mars-encounter, the sail functions parachute-like in Mars s outer atmosphere to accomplish aerocapture. On-board thrusters or the sail maneuver the spacecraft into an orbit with periapsis near Mars and apoapsis near Phobos. The orbit is circularized for Phobos-rendezvous; surface samples are collected. The sail then raises the orbit for Deimos-rendezvous and sample collection. The sail next places the spacecraft on an Earth-bound Hohmann-transfer ellipse. During Earth-encounter, the sail accomplishes Earth-aerocapture or partially decelerates the sample container for entry into Earth s atmosphere. Mission mass budget is about 218 grams and; mission duration is <5 years.

KSC-108-75PC-0388

NASA Image and Video Library

1975-07-15

CAPE CANAVERAL, Fla. – The Apollo Soyuz Test Project Saturn IB launch vehicle thundered away from KSC’s Launch Complex 39B at 3:50 p.m. today. Aboard the Apollo Command Module were ASTP Astronauts Thomas Stafford, Vance Brand and Donald Slayton. The astronauts will rendezvous and dock with a Soyuz spacecraft, launched this morning from the Baikonur launch facility in the Soviet Union, carrying Soviet cosmonauts Aleksey Leonov and Valeriy Kubasov. The first international crewed spaceflight was a joint U.S.-U.S.S.R. rendezvous and docking mission. The Apollo-Soyuz Test Project, or ASTP, took its name from the spacecraft employed: the American Apollo and the Soviet Soyuz. The three-man Apollo crew lifted off from Kennedy Space Center aboard a Saturn IB rocket on July 15, 1975, to link up with the Soyuz that had launched a few hours earlier. A cylindrical docking module served as an airlock between the two spacecraft for transfer of the crew members. Photo credit: NASA

KSC-04pd1824

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, Corky Philyaw (left) and Edgar Suarez (right) prepare the flight battery for installation on the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft (far left). DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. It is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA's Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station. DART will be launched from an Orbital Sciences Pegasus XL rocket no earlier than Oct. 26.

KSC-04pd1817

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers prepare the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft for launch. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Orbital Sciences Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

A Comparison Between Orion Automated and Space Shuttle Rendezvous Techniques

NASA Technical Reports Server (NTRS)

Ruiz, Jose O,; Hart, Jeremy

2010-01-01

The Orion spacecraft will replace the space shuttle and will be the first human spacecraft since the Apollo program to leave low earth orbit. This vehicle will serve as the cornerstone of a complete space transportation system with a myriad of mission requirements necessitating rendezvous to multiple vehicles in earth orbit, around the moon and eventually beyond . These goals will require a complex and robust vehicle that is, significantly different from both the space shuttle and the command module of the Apollo program. Historically, orbit operations have been accomplished with heavy reliance on ground support and manual crew reconfiguration and monitoring. One major difference with Orion is that automation will be incorporated as a key element of the man-vehicle system. The automated system will consist of software devoted to transitioning between events based on a master timeline. This effectively adds a layer of high level sequencing that moves control of the vehicle from one phase to the next. This type of automated control is not entirely new to spacecraft since the shuttle uses a version of this during ascent and entry operations. During shuttle orbit operations however many of the software modes and hardware switches must be manually configured through the use of printed procedures and instructions voiced from the ground. The goal of the automation scheme on Orion is to extend high level automation to all flight phases. The move towards automation represents a large shift from current space shuttle operations, and so these new systems will be adopted gradually via various safeguards. These include features such as authority-to-proceed, manual down modes, and functional inhibits. This paper describes the contrast between the manual and ground approach of the space shuttle and the proposed automation of the Orion vehicle. I will introduce typical orbit operations that are common to all rendezvous missions and go on to describe the current Orion automation architecture and contrast it with shuttle rendezvous techniques and circumstances. The shuttle rendezvous profile is timed to take approximately 3 days from orbit insertion to docking at the International Space Station (ISS). This process can be divided into 3 phases: far-field, mid-field and proximity operations. The far-field stage is characterized as the most quiescent phase. The spacecraft is usually too far to navigate using relative sensors and uses the Inertial Measurement Units (IMU s) to numerically solve for its position. The maneuvers are infrequent, roughly twice per day, and are larger than other burns in the profile. The shuttle uses this opportunity to take extensive ground based radar updates and keep high fidelity orbit states on the ground. This state is then periodically uplinked to the shuttle computers. The targeting solutions for burn maneuvers are also computed on the ground and uplinked. During the burn the crew is responsible for setting the shuttle attitude and configuring the propulsion system for ignition. Again this entire process is manually driven by both crew and ground activity. The only automatic processes that occur are associated with the real-time execution of the burn. The Orion automated functionality will seek to relieve the workload of both the crew and ground during this phase

Path scheduling for multiple mobile actors in wireless sensor network

NASA Astrophysics Data System (ADS)

Trapasiya, Samir D.; Soni, Himanshu B.

2017-05-01

In wireless sensor network (WSN), energy is the main constraint. In this work we have addressed this issue for single as well as multiple mobile sensor actor network. In this work, we have proposed Rendezvous Point Selection Scheme (RPSS) in which Rendezvous Nodes are selected by set covering problem approach and from that, Rendezvous Points are selected in a way to reduce the tour length. The mobile actors tour is scheduled to pass through those Rendezvous Points as per Travelling Salesman Problem (TSP). We have also proposed novel rendezvous node rotation scheme for fair utilisation of all the nodes. We have compared RPSS with Stationery Actor scheme as well as RD-VT, RD-VT-SMT and WRP-SMT for performance metrics like energy consumption, network lifetime, route length and found the better outcome in all the cases for single actor. We have also applied RPSS for multiple mobile actor case like Multi-Actor Single Depot (MASD) termination and Multi-Actor Multiple Depot (MAMD) termination and observed by extensive simulation that MAMD saves the network energy in optimised way and enhance network lifetime compared to all other schemes.

A study of unmanned mission opportunities to comets and asteroids

NASA Technical Reports Server (NTRS)

Mann, F. I.; Horsewood, J. L.; Bjorkman, W.

1974-01-01

Several unmanned multiple-target mission opportunities to comets and asteroids were studied. The targets investigated include Grigg-Skjellerup, Giacobini-Zinner, Tuttle-Giacobini-Kresak, Borrelly, Halley, Schaumasse, Geographos, Eros, Icarus, and Toro, and the trajectories consist of purely ballistic flight, except that powered swingbys and deep space burns are employed when necessary. Optimum solar electric rendezvous trajectories to the comets Giacobini-Zinner/85, Borrelly/87, and Temple (2)/83 and /88 employing the 8.67 kw Sert III spacecraft modified for interplanetary flight were also investigated. The problem of optimizing electric propulsion heliocentric trajectories, including the effects of geocentric launch asymptote declination on launch vehicle performance capability, was formulated, and a solution developed using variational calculus techniques. Improvements were made to the HILTOP trajectory optimization computer program. An error analysis of high-thrust maneuvers involving spin-stabilized spacecraft was developed and applied to a synchronous meteorological satellite mission.

OSIRIS-REx Orbit Determination Covariance Studies at Bennu

NASA Technical Reports Server (NTRS)

Antreasian, P. G.; Moreau, M.; Jackman, C.; Williams, K.; Page, B.; Leonard, J. M.

2016-01-01

The Origins Spectral Interpretation Resource Identification Security Regolith Explorer (OSIRIS-REx) mission is a NASA New Frontiers mission launching in 2016 to rendezvous with the small, Earth-crossing asteroid (101955) Bennu in late 2018, and ultimately return a sample of regolith to Earth. Approximately 3 months before the encounter with Bennu, the asteroid finally becomes detectable in the narrow field PolyCam imager. The spacecraft's rendezvous with Bennu begins with a series of four Asteroid Approach Maneuvers, which slow the spacecraft's speed relative to Bennu beginning two and a half months prior to closest approach, ultimately delivering the spacecraft to a point 18 km from Bennu on Nov 18, 2018. An extensive campaign of proximity operations activities to characterize the properties of Bennu and select a suitable sample site will follow. This paper will discuss the challenges of navigating near a small 500-m diameter asteroid. The navigation at close proximity is dependent on the accurate mathematical model or digital terrain map of the asteroids shape. Predictions of the spacecraft state are very sensitive to spacecraft small forces, solar radiation pressure, and mis-modeling of Bennu's gravity field. Uncertainties in the physical parameters of the central body Bennu create additional challenges. The navigation errors are discussed and their impact on science planning will be presented.

OSIRIS-REx Orbit Determination Covariance Studies at Bennu

NASA Technical Reports Server (NTRS)

Antreasian, P. G.; Moreau, M.; Jackman, C.; Williams, K.; Page, B.; Leonard, J. M.

2016-01-01

The Origins Spectral Interpretation Resource Identification Security Regolith Explorer (OSIRIS-REx) mission is a NASA New Frontiers mission launching in 2016 to rendezvous with the small, Earth-crossing asteroid (101955) Bennu in late 2018, ultimately returning a sample of regolith to Earth. Approximately three months before the encounter with Bennu, the asteroid becomes detectable in the narrow field PolyCam imager. The spacecraft's rendezvous with Bennu begins with a series of four Asteroid Approach Maneuvers, slowing the spacecraft's speed relative to Bennu beginning two and a half months prior to closest approach, ultimately delivering the spacecraft to a point 18 km from Bennu in Nov, 2018. An extensive campaign of proximity operations activities to characterize the properties of Bennu and select a suitable sample site will follow. This paper will discuss the challenges of navigating near a small 500-m diameter asteroid. The navigation at close proximity is dependent on the accurate mathematical model or digital terrain map of the asteroid's shape. Predictions of the spacecraft state are very sensitive to spacecraft small forces, solar radiation pressure, and mis-modeling of Bennu's gravity field. Uncertainties in the physical parameters of the central body Bennu create additional challenges. The navigation errors are discussed and their impact on science planning will be presented.

GEMINI-TITAN (GT)-9 - EARTH-SKY - AUGMENTED TARGET DOCKING ADAPTER (ATDA) - MSC

NASA Image and Video Library

1966-06-06

S66-37923 (3 June 1966) --- The Augmented Target Docking Adapter (ATDA) as seen from the Gemini-9 spacecraft during one of their three rendezvous in space. The ATDA and Gemini-9 spacecraft are 66.5 feet apart. Failure of the docking adapter protective cover to fully separate on the ATDA prevented the docking of the two spacecraft. The ATDA was described by the Gemini-9 crew as an "angry alligator." Photo credit: NASA

GEMINI-9 - EARTH SKY - ATDA

NASA Image and Video Library

1966-06-06

S66-37943 (3 June 1966) --- The Augmented Target Docking Adapter is photographed against the background of the blackness of space from the Gemini-9 spacecraft during one of their three rendezvous in space. The ATDA and Gemini-9 spacecraft are 71.5 feet apart. Failure of the docking adapter protective cover to fully separate on the ATDA prevented the docking of the two spacecraft. The ATDA was described by the Gemini-9 crew as an ?Angry Alligator.? Photo credit: NASA

Pose Measurement Performance of the Argon Relative Navigation Sensor Suite in Simulated Flight Conditions

NASA Technical Reports Server (NTRS)

Galante, Joseph M.; Eepoel, John Van; Strube, Matt; Gill, Nat; Gonzalez, Marcelo; Hyslop, Andrew; Patrick, Bryan

2012-01-01

Argon is a flight-ready sensor suite with two visual cameras, a flash LIDAR, an on- board flight computer, and associated electronics. Argon was designed to provide sensing capabilities for relative navigation during proximity, rendezvous, and docking operations between spacecraft. A rigorous ground test campaign assessed the performance capability of the Argon navigation suite to measure the relative pose of high-fidelity satellite mock-ups during a variety of simulated rendezvous and proximity maneuvers facilitated by robot manipulators in a variety of lighting conditions representative of the orbital environment. A brief description of the Argon suite and test setup are given as well as an analysis of the performance of the system in simulated proximity and rendezvous operations.

Reduced cost mission design using surrogate models

NASA Astrophysics Data System (ADS)

Feldhacker, Juliana D.; Jones, Brandon A.; Doostan, Alireza; Hampton, Jerrad

2016-01-01

This paper uses surrogate models to reduce the computational cost associated with spacecraft mission design in three-body dynamical systems. Sampling-based least squares regression is used to project the system response onto a set of orthogonal bases, providing a representation of the ΔV required for rendezvous as a reduced-order surrogate model. Models are presented for mid-field rendezvous of spacecraft in orbits in the Earth-Moon circular restricted three-body problem, including a halo orbit about the Earth-Moon L2 libration point (EML-2) and a distant retrograde orbit (DRO) about the Moon. In each case, the initial position of the spacecraft, the time of flight, and the separation between the chaser and the target vehicles are all considered as design inputs. The results show that sample sizes on the order of 102 are sufficient to produce accurate surrogates, with RMS errors reaching 0.2 m/s for the halo orbit and falling below 0.01 m/s for the DRO. A single function call to the resulting surrogate is up to two orders of magnitude faster than computing the same solution using full fidelity propagators. The expansion coefficients solved for in the surrogates are then used to conduct a global sensitivity analysis of the ΔV on each of the input parameters, which identifies the separation between the spacecraft as the primary contributor to the ΔV cost. Finally, the models are demonstrated to be useful for cheap evaluation of the cost function in constrained optimization problems seeking to minimize the ΔV required for rendezvous. These surrogate models show significant advantages for mission design in three-body systems, in terms of both computational cost and capabilities, over traditional Monte Carlo methods.

Automated Rendezvous and Capture System Development and Simulation for NASA

NASA Technical Reports Server (NTRS)

Roe, Fred D.; Howard, Richard T.; Murphy, Leslie

2004-01-01

The United States does not have an Automated Rendezvous and Capture Docking (AR&C) capability and is reliant on manned control for rendezvous and docking of orbiting spacecraft. T h i s reliance on the labor intensive manned interface for control of rendezvous and docking vehicles has a significant impact on the cost of the operation of the International Space Station (ISS) and precludes the use of any U.S. expendable launch capabilities for Space Station resupply. The Marshall Space Flight Center (MSFC) has conducted pioneering research in the development of an automated rendezvous and capture (or docking) (AR&C) system for U.S. space vehicles. This A M C system was tested extensively using hardware-in-the-loop simulations in the Flight Robotics Laboratory, and a rendezvous sensor, the Video Guidance Sensor was developed and successfully flown on the Space Shuttle on flights STS-87 and STS-95, proving the concept of a video- based sensor. Further developments in sensor technology and vehicle and target configuration have lead to continued improvements and changes in AR&C system development and simulation. A new Advanced Video Guidance Sensor (AVGS) with target will be utilized as the primary navigation sensor on the Demonstration of Autonomous Rendezvous Technologies (DART) flight experiment in 2004. Realtime closed-loop simulations will be performed to validate the improved AR&C systems prior to flight.

Enchanted rendezvous: John C. Houbolt and the genesis of the lunar-orbit rendezvous concept

NASA Technical Reports Server (NTRS)

Hansen, James R.

1995-01-01

This is the fourth publication of the 'Monographs in Aerospace History' series, prepared by the NASA History Office. These publications are intended to be tightly focused in terms of subject, relatively short in length, and reproduced to allow timely and broad dissemination to researchers in aerospace history. This publication details the arguments of John C. Houbolt, an engineer at the Langley Research Center in Hampton, Virginia, in his 1961-1962 campaign to support the lunar-orbit rendezvous (LOR). The LOR was eventually selected during Project Apollo as the method of flying to the Moon, landing on the surface, and returning to Earth. The LOR opted to send the entire lunar spacecraft up in one launch, enter into the lunar orbit, and dispatch a small lander to the lunar surface. It was the simplest of the various methods, both in terms of development and operational costs, but it was risky. There was no room for error or the crew could not get home; and the more difficult maneuvers had to be done when the spacecraft was committed to a circumlunar flight. Houbolt was one of the most vocal people supporting the LOR.

KSC-04pd1819

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers help guide the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft onto the mobile stand below. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Orbital Sciences Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

Robust model predictive control for multi-step short range spacecraft rendezvous

NASA Astrophysics Data System (ADS)

Zhu, Shuyi; Sun, Ran; Wang, Jiaolong; Wang, Jihe; Shao, Xiaowei

2018-07-01

This work presents a robust model predictive control (MPC) approach for the multi-step short range spacecraft rendezvous problem. During the specific short range phase concerned, the chaser is supposed to be initially outside the line-of-sight (LOS) cone. Therefore, the rendezvous process naturally includes two steps: the first step is to transfer the chaser into the LOS cone and the second step is to transfer the chaser into the aimed region with its motion confined within the LOS cone. A novel MPC framework named after Mixed MPC (M-MPC) is proposed, which is the combination of the Variable-Horizon MPC (VH-MPC) framework and the Fixed-Instant MPC (FI-MPC) framework. The M-MPC framework enables the optimization for the two steps to be implemented jointly rather than to be separated factitiously, and its computation workload is acceptable for the usually low-power processors onboard spacecraft. Then considering that disturbances including modeling error, sensor noise and thrust uncertainty may induce undesired constraint violations, a robust technique is developed and it is attached to the above M-MPC framework to form a robust M-MPC approach. The robust technique is based on the chance-constrained idea, which ensures that constraints can be satisfied with a prescribed probability. It improves the robust technique proposed by Gavilan et al., because it eliminates the unnecessary conservativeness by explicitly incorporating known statistical properties of the navigation uncertainty. The efficacy of the robust M-MPC approach is shown in a simulation study.

An Application of Linear Covariance Analysis to the Design of Responsive Near-Rendezvous Missions

DTIC Science & Technology

2007-06-01

accurately before making large ma- neuvers. A fifth type of error is maneuver knowledge error (MKER). This error accounts for how well a spacecraft is able...utilized due in a large part to the cost of designing and launching spacecraft , in a market where currently there are not many options for launching...is then ordered to fire its thrusters to increase its orbital altitude to 800 km. Before the maneuver the spacecraft is moving with some velocity, V

Flight data results of estimate fusion for spacecraft rendezvous navigation from shuttle mission STS-69

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell; Bishop, Robert H.

1996-01-01

A recently developed rendezvous navigation fusion filter that optimally exploits existing distributed filters for rendezvous and GPS navigation to achieve the relative and inertial state accuracies of both in a global solution is utilized here to process actual flight data. Space Shuttle Mission STS-69 was the first mission to date which gathered data from both the rendezvous and Global Positioning System filters allowing, for the first time, a test of the fusion algorithm with real flight data. Furthermore, a precise best estimate of trajectory is available for portions of STS-69, making possible a check on the performance of the fusion filter. In order to successfully carry out this experiment with flight data, two extensions to the existing scheme were necessary: a fusion edit test based on differences between the filter state vectors, and an underweighting scheme to accommodate the suboptimal perfect target assumption made by the Shuttle rendezvous filter. With these innovations, the flight data was successfully fused from playbacks of downlinked and/or recorded measurement data through ground analysis versions of the Shuttle rendezvous filter and a GPS filter developed for another experiment. The fusion results agree with the best estimate of trajectory at approximately the levels of uncertainty expected from the fusion filter's covariance matrix.

Planetary mission requirements, technology and design considerations for a solar electric propulsion stage

NASA Technical Reports Server (NTRS)

Cork, M. J.; Hastrup, R. C.; Menard, W. A.; Olson, R. N.

1979-01-01

High energy planetary missions such as comet rendezvous, Saturn orbiter and asteroid rendezvous require development of a Solar Electric Propulsion Stage (SEPS) for augmentation of the Shuttle-IUS. Performance and functional requirements placed on the SEPS are presented. These requirements will be used in evolution of the SEPS design, which must be highly interactive with both the spacecraft and the mission design. Previous design studies have identified critical SEPS technology areas and some specific design solutions which are also presented in the paper.

Dual RF Astrodynamic GPS Orbital Navigator Satellite

NASA Technical Reports Server (NTRS)

Kanipe, David B.; Provence, Robert Steve; Straube, Timothy M.; Reed, Helen; Bishop, Robert; Lightsey, Glenn

2009-01-01

Dual RF Astrodynamic GPS Orbital Navigator Satellite (DRAGONSat) will demonstrate autonomous rendezvous and docking (ARD) in low Earth orbit (LEO) and gather flight data with a global positioning system (GPS) receiver strictly designed for space applications. ARD is the capability of two independent spacecraft to rendezvous in orbit and dock without crew intervention. DRAGONSat consists of two picosatellites (one built by the University of Texas and one built by Texas A and M University) and the Space Shuttle Payload Launcher (SSPL); this project will ultimately demonstrate ARD in LEO.

Earth Science

NASA Image and Video Library

1996-01-31

The Near Earth Asteroid Rendezvous (NEAR) spacecraft embarks on a journey that will culminate in a close encounter with an asteroid. The launch of NEAR inaugurates NASA's irnovative Discovery program of small-scale planetary missions with rapid, lower-cost development cycles and focused science objectives. NEAR will rendezvous in 1999 with the asteroid 433 Eros to begin the first long-term, close-up look at an asteroid's surface composition and physical properties. NEAR's science payload includes an x-ray/gamma ray spectrometer, an near-infrared spectrograph, a laser rangefinder, a magnetometer, a radio science experiment and a multi-spectral imager.

Re-rendezvous and approach of Progress 33P

NASA Image and Video Library

2009-07-12

ISS020-E-018056 (12 July 2009) --- An unpiloted ISS Progress 33 cargo craft approaches the International Space Station. On June 30, the Progress undocked from the station and was commanded into a parking orbit for its re-rendezvous with the ISS on July 12, approaching to within 10-15 meters of the Zvezda Service Module to test new automated rendezvous equipment mounted on Zvezda during a pair of spacewalks earlier this month by Gennady Padalka and Mike Barratt that will be used to guide the new Mini-Research Module-2 (MRM2) to an unpiloted docking to the zenith port of Zvezda later this year. MRM2 will serve as a new docking port for Russian spacecraft and an additional airlock for spacewalks conducted out of the Russian segment.

Antenna and propagation studies for spacecraft systems: Addendum to preflight SL-1/SL-3 Skylab VHF ranging coverage (early and late TPI)

NASA Technical Reports Server (NTRS)

Eisenhauer, D. R.; James, D. A.

1973-01-01

A preflight assessment is presented of the expected Skylab VHF ranging coverage for the rendezvous portion of the SL-1/SL-3 mission, assuming a 28 July 1973 launch date, for the alternative trajectory cases characterized by either an early TPI or a late TPI. In this assessment early TPI and late TPI are used to indicate a TPI maneuver occurring 10 minutes prior to or after the nominally scheduled TPI maneuver, respectively. The Saturn workshop (SWS) maintains a solar inertial (SI) attitude throughout rendezvous for both trajectory cases. The results summarized concern VHF ranging function performance during that period most likely to be affected by off-nominal TPI conditions, i.e., NSR (5:56 g.e.t.) to station keeping. Curves are presented which show the variation in received power levels on both spacecraft-to-spacecraft links from about 100 n.mi. range to CSM and SWS station keeping. Appropriate threshold levels are shown on these received power curves to indicate zero circuit margins for the ranging function.

Video guidance, landing, and imaging systems

NASA Technical Reports Server (NTRS)

Schappell, R. T.; Knickerbocker, R. L.; Tietz, J. C.; Grant, C.; Rice, R. B.; Moog, R. D.

1975-01-01

The adaptive potential of video guidance technology for earth orbital and interplanetary missions was explored. The application of video acquisition, pointing, tracking, and navigation technology was considered to three primary missions: planetary landing, earth resources satellite, and spacecraft rendezvous and docking. It was found that an imaging system can be mechanized to provide a spacecraft or satellite with a considerable amount of adaptability with respect to its environment. It also provides a level of autonomy essential to many future missions and enhances their data gathering ability. The feasibility of an autonomous video guidance system capable of observing a planetary surface during terminal descent and selecting the most acceptable landing site was successfully demonstrated in the laboratory. The techniques developed for acquisition, pointing, and tracking show promise for recognizing and tracking coastlines, rivers, and other constituents of interest. Routines were written and checked for rendezvous, docking, and station-keeping functions.

KSC-04pd1826

NASA Image and Video Library

2004-09-02

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft (right) is ready for mating with the upper stage (foreground) in preparation for launch on the Orbital Sciences Pegasus XL. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1830

NASA Image and Video Library

2004-09-03

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers maneuver the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft and mated upper stage toward the second stage at right in preparation or launch aboard the Orbital Sciences Pegasus XL launch vehicle. Pegasus will launch DART into a circular polar orbit of approximately 475 miles. Built for NASA by Orbital Sciences Corporation, DART was designed as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1827

NASA Image and Video Library

2004-09-02

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers maneuver the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft, suspended by a crane, over the upper stage in preparation for launch on the Orbital Sciences Pegasus XL. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. Built for NASA by Orbital Sciences Corporation, DART was designed as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1820

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft (in background) has been rotated from vertical to horizontal and is ready for mating with the upper stage (foreground). DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Orbital Sciences Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1823

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers begin closing the gap between the second and third stages of the Orbital Sciences Pegasus XL launch vehicle that will launch the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA's Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1828

NASA Image and Video Library

2004-09-03

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers maneuver the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft and mated upper stage toward the second stage behind them in preparation or launch aboard the Orbital Sciences Pegasus XL launch vehicle. Pegasus will launch DART into a circular polar orbit of approximately 475 miles. Built for NASA by Orbital Sciences Corporation, DART was designed as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1816

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, a worker prepares the second and third stages of the Orbital Sciences Pegasus XL launch vehicle for mating. The Pegasus XL will launch the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1825

NASA Image and Video Library

2004-09-02

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft (right) is ready for mating with the upper stage (behind it) in preparation for launch on the Orbital Sciences Pegasus XL. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1818

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers stand by while an overhead crane moves the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft onto the mobile stand at right. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Orbital Sciences Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1821

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft is ready for mating with the upper stage of the Orbital Sciences Pegasus XL behind it (right). DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1822

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers begin mating the second and third stages of the Orbital Sciences Pegasus XL launch vehicle that will launch the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA's Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1829

NASA Image and Video Library

2004-09-03

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft (foreground) is ready to be mated to second and third stages in preparation for the launch aboard the Orbital Sciences Pegasus XL launch vehicle. Pegasus will launch DART into a circular polar orbit of approximately 475 miles. Built for NASA by Orbital Sciences Corporation, DART was designed as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA’s Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04pd1815

NASA Image and Video Library

2004-09-01

KENNEDY SPACE CENTER, FLA. - At Vandenberg Air Force Base in California, workers prepare to mate the second and third stages of the Orbital Sciences Pegasus XL launch vehicle that will launch the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASA's Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

Autonomous rendezvous and docking: A commercial approach to on-orbit technology validation

NASA Technical Reports Server (NTRS)

Tchoryk, Peter, Jr.; Whitten, Raymond P.

1991-01-01

SpARC, in conjunction with its corporate affiliates, is planning an on-orbit validation of autonomous rendezvous and docking (ARD) technology. The emphasis in this program is to utilize existing technology and commercially available components wherever possible. The primary subsystems to be validated by this demonstration include GPS receivers for navigation, a video-based sensor for proximity operations, a fluid connector mechanism to demonstrate fluid resupply capability, and a compliant, single-point docking mechanism. The focus for this initial experiment will be ELV based and will make use of two residual Commercial Experiment Transporter (COMET) service modules. The first COMET spacecraft will be launched in late 1992 and will serve as the target vehicle. After the second COMET spacecraft has been launched in late 1994, the ARD demonstration will take place. The service module from the second COMET will serve as the chase vehicle.

KSC-04PD-1818

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. At Vandenberg Air Force Base in California, workers stand by while an overhead crane moves the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft onto the mobile stand at right. DART was designed and built for NASA by Orbital Sciences Corporation as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. The Orbital Sciences Pegasus XL will launch DART into a circular polar orbit of approximately 475 miles. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASAs Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

KSC-04PD-1830

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. At Vandenberg Air Force Base in California, workers maneuver the Demonstration of Autonomous Rendezvous Technology (DART) spacecraft and mated upper stage toward the second stage at right in preparation or launch aboard the Orbital Sciences Pegasus XL launch vehicle. Pegasus will launch DART into a circular polar orbit of approximately 475 miles. Built for NASA by Orbital Sciences Corporation, DART was designed as an advanced flight demonstrator to locate and maneuver near an orbiting satellite. DART weighs about 800 pounds and is nearly 6 feet long and 3 feet in diameter. DART is designed to demonstrate technologies required for a spacecraft to locate and rendezvous, or maneuver close to, other craft in space. Results from the DART mission will aid in the development of NASAs Crew Exploration Vehicle and will also assist in vehicle development for crew transfer and crew rescue capability to and from the International Space Station.

Antares Post Launch Press Conference

NASA Image and Video Library

2013-09-18

Frank Culbertson, executive vice president, Orbital Sciences Corporation, talks during a press conference held after the successful launch of the Antares rocket, with the Cygnus cargo spacecraft aboard, Wednesday, Sept. 18, 2013, NASA Wallops Flight Facility, Virginia. Cygnus is on its way to rendezvous with the space station. The spacecraft will deliver about 1,300 pounds (589 kilograms) of cargo, including food and clothing, to the Expedition 37 crew. Photo Credit: (NASA/Bill Ingalls)

Antares Post Launch Press Conference

NASA Image and Video Library

2013-09-18

Robert Lightfoot, associate administrator, NASA, talks during a press conference held after the successful launch of the Antares rocket, with the Cygnus cargo spacecraft aboard, Wednesday, Sept. 18, 2013, NASA Wallops Flight Facility, Virginia. Cygnus is on its way to rendezvous with the space station. The spacecraft will deliver about 1,300 pounds (589 kilograms) of cargo, including food and clothing, to the Expedition 37 crew. Photo Credit: (NASA/Bill Ingalls)

SPHERES experiment session

NASA Image and Video Library

2007-03-24

ISS014-E-17880 (24 March 2007) --- This medium close-up view shows three bowling-ball-sized free-flying satellites called Synchronized Position Hold, Engage, Reorient, Experimental Satellites (SPHERES) in the Destiny laboratory of the International Space Station. SPHERES were designed to test control algorithms for spacecraft by performing autonomous rendezvous and docking maneuvers inside the station. The results are important for multi-body control and in designing constellation and array spacecraft configurations.

Navigation for Rendezvous and Orbit Missions to Small Solar-System Bodies

NASA Technical Reports Server (NTRS)

Helfrich, C. E.; Scheeres, D. J.; Williams, B. G.; Bollman, W. E.; Davis, R. P.; Synnott, S. P.; Yeomans, D. K.

1994-01-01

All previous spacecraft encounters with small solar-system bodies, such as asteroids and comets, have been flybys (e.g. Galileo's flybys of the asteroids Gaspra and Ida). Several future projects plan to build on the flyby experience and progress to the next level with rendezvous and orbit missions to small bodies. This presents several new issues and challenges for navigation which have never been considered before. This paper addresses these challenges by characterizing the different phases of a small body rendezvous and by describing the navigation requirements and goals of each phase. Prior to the encounter with the small body, improvements to its ephemeris and initial estimates of its physical parameters, e.g. size, shape, mass, rotation rate, rotation pole, and possibly outgassing, are made as accurately as ground-based measurements allow. This characterization can take place over years...

GN/C translation and rotation control parameters for AR/C (category 2)

NASA Technical Reports Server (NTRS)

Henderson, David M.

1991-01-01

Detailed analysis of the Automatic Rendezvous and Capture problem indicate a need for three different regions of mathematical description for the GN&C algorithms: (1) multi-vehicle orbital mechanics to the rendezvous interface point, i.e., within 100 n.; (2) relative motion solutions (such as Clohessy-Wiltshire type) from the far-field to the near-field interface, i.e., within 1 nm; and (3) close proximity motion, the nearfield motion where the relative differences in the gravitational and orbit inertial accelerations can be neglected from the equations of motion. This paper defines the reference coordinate frames and control parameters necessary to model the relative motion and attitude of spacecraft in the close proximity of another space system (Region 2 and 3) during the Automatic Rendezvous and Capture phase of an orbit operation.

Automated Rendezvous and Capture System Development and Simulation for NASA

NASA Technical Reports Server (NTRS)

Roe, Fred D.; Howard, Richard T.; Murphy, Leslie

2004-01-01

The United States does not have an Automated Rendezvous and Capture/Docking (AR and C) capability and is reliant on manned control for rendezvous and docking of orbiting spacecraft. This reliance on the labor intensive manned interface for control of rendezvous and docking vehicles has a significant impact on the cost of the operation of the International Space Station (ISS) and precludes the use of any U.S. expendable launch capabilities for Space Station resupply. The Soviets have the capability to autonomously dock in space, but their system produces a hard docking with excessive force and contact velocity. Automated Rendezvous and Capture/Docking has been identified as a key enabling technology for the Space Launch Initiative (SLI) Program, DARPA Orbital Express and other DOD Programs. The development and implementation of an AR&C capability can significantly enhance system flexibility, improve safety, and lower the cost of maintaining, supplying, and operating the International Space Station. The Marshall Space Flight Center (MSFC) has conducted pioneering research in the development of an automated rendezvous and capture (or docking) (AR and C) system for U.S. space vehicles. This AR&C system was tested extensively using hardware-in-the-loop simulations in the Flight Robotics Laboratory, and a rendezvous sensor, the Video Guidance Sensor was developed and successfully flown on the Space Shuttle on flights STS-87 and STS-95, proving the concept of a video- based sensor. Further developments in sensor technology and vehicle and target configuration have lead to continued improvements and changes in AR&C system development and simulation. A new Advanced Video Guidance Sensor (AVGS) with target will be utilized on the Demonstration of Autonomous Rendezvous Technologies (DART) flight experiment in 2004.

ADRC for spacecraft attitude and position synchronization in libration point orbits

NASA Astrophysics Data System (ADS)

Gao, Chen; Yuan, Jianping; Zhao, Yakun

2018-04-01

This paper addresses the problem of spacecraft attitude and position synchronization in libration point orbits between a leader and a follower. Using dual quaternion, the dimensionless relative coupled dynamical model is derived considering computation efficiency and accuracy. Then a model-independent dimensionless cascade pose-feedback active disturbance rejection controller is designed to spacecraft attitude and position tracking control problems considering parameter uncertainties and external disturbances. Numerical simulations for the final approach phase in spacecraft rendezvous and docking and formation flying are done, and the results show high-precision tracking errors and satisfactory convergent rates under bounded control torque and force which validate the proposed approach.

Adaptive relative pose control of spacecraft with model couplings and uncertainties

NASA Astrophysics Data System (ADS)

Sun, Liang; Zheng, Zewei

2018-02-01

The spacecraft pose tracking control problem for an uncertain pursuer approaching to a space target is researched in this paper. After modeling the nonlinearly coupled dynamics for relative translational and rotational motions between two spacecraft, position tracking and attitude synchronization controllers are developed independently by using a robust adaptive control approach. The unknown kinematic couplings, parametric uncertainties, and bounded external disturbances are handled with adaptive updating laws. It is proved via Lyapunov method that the pose tracking errors converge to zero asymptotically. Spacecraft close-range rendezvous and proximity operations are introduced as an example to validate the effectiveness of the proposed control approach.

Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation

NASA Technical Reports Server (NTRS)

Miller, J. K.; Antreasian, P. J.; Georgini, J.; Owen, W. M.; Williams, B. G.; Yeomans, D. K.

1995-01-01

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.

ISS Expedition 18 Synchronized Position Hold,Engage,Reorient,Experimental Satellites (SPHERES)

NASA Image and Video Library

2008-10-26

ISS018-E-005214 (26 Oct. 2008) --- This close-up view shows three bowling-ball-sized free-flying satellites called Synchronized Position Hold, Engage, Reorient, Experimental Satellites (SPHERES) in the Destiny laboratory of the International Space Station. SPHERES were designed to test control algorithms for spacecraft by performing autonomous rendezvous and docking maneuvers inside the station. The results are important for multi-body control and in designing constellation and array spacecraft configurations.

Proximity Operations for Space Situational Awareness Spacecraft Rendezvous and Maneuvering using Numerical Simulations and Fuzzy Logic

NASA Astrophysics Data System (ADS)

Carrico, T.; Langster, T.; Carrico, J.; Alfano, S.; Loucks, M.; Vallado, D.

The authors present several spacecraft rendezvous and close proximity maneuvering techniques modeled with a high-precision numerical integrator using full force models and closed loop control with a Fuzzy Logic intelligent controller to command the engines. The authors document and compare the maneuvers, fuel use, and other parameters. This paper presents an innovative application of an existing capability to design, simulate and analyze proximity maneuvers; already in use for operational satellites performing other maneuvers. The system has been extended to demonstrate the capability to develop closed loop control laws to maneuver spacecraft in close proximity to another, including stand-off, docking, lunar landing and other operations applicable to space situational awareness, space based surveillance, and operational satellite modeling. The fully integrated end-to-end trajectory ephemerides are available from the authors in electronic ASCII text by request. The benefits of this system include: A realistic physics-based simulation for the development and validation of control laws A collaborative engineering environment for the design, development and tuning of spacecraft law parameters, sizing actuators (i.e., rocket engines), and sensor suite selection. An accurate simulation and visualization to communicate the complexity, criticality, and risk of spacecraft operations. A precise mathematical environment for research and development of future spacecraft maneuvering engineering tasks, operational planning and forensic analysis. A closed loop, knowledge-based control example for proximity operations. This proximity operations modeling and simulation environment will provide a valuable adjunct to programs in military space control, space situational awareness and civil space exploration engineering and decision making processes.

Gossip-based solutions for discrete rendezvous in populations of communicating agents.

PubMed

Hollander, Christopher D; Wu, Annie S

2014-01-01

The objective of the rendezvous problem is to construct a method that enables a population of agents to agree on a spatial (and possibly temporal) meeting location. We introduce the buffered gossip algorithm as a general solution to the rendezvous problem in a discrete domain with direct communication between decentralized agents. We compare the performance of the buffered gossip algorithm against the well known uniform gossip algorithm. We believe that a buffered solution is preferable to an unbuffered solution, such as the uniform gossip algorithm, because the use of a buffer allows an agent to use multiple information sources when determining its desired rendezvous point, and that access to multiple information sources may improve agent decision making by reinforcing or contradicting an initial choice. To show that the buffered gossip algorithm is an actual solution for the rendezvous problem, we construct a theoretical proof of convergence and derive the conditions under which the buffered gossip algorithm is guaranteed to produce a consensus on rendezvous location. We use these results to verify that the uniform gossip algorithm also solves the rendezvous problem. We then use a multi-agent simulation to conduct a series of simulation experiments to compare the performance between the buffered and uniform gossip algorithms. Our results suggest that the buffered gossip algorithm can solve the rendezvous problem faster than the uniform gossip algorithm; however, the relative performance between these two solutions depends on the specific constraints of the problem and the parameters of the buffered gossip algorithm.

Gossip-Based Solutions for Discrete Rendezvous in Populations of Communicating Agents

PubMed Central

Hollander, Christopher D.; Wu, Annie S.

2014-01-01

The objective of the rendezvous problem is to construct a method that enables a population of agents to agree on a spatial (and possibly temporal) meeting location. We introduce the buffered gossip algorithm as a general solution to the rendezvous problem in a discrete domain with direct communication between decentralized agents. We compare the performance of the buffered gossip algorithm against the well known uniform gossip algorithm. We believe that a buffered solution is preferable to an unbuffered solution, such as the uniform gossip algorithm, because the use of a buffer allows an agent to use multiple information sources when determining its desired rendezvous point, and that access to multiple information sources may improve agent decision making by reinforcing or contradicting an initial choice. To show that the buffered gossip algorithm is an actual solution for the rendezvous problem, we construct a theoretical proof of convergence and derive the conditions under which the buffered gossip algorithm is guaranteed to produce a consensus on rendezvous location. We use these results to verify that the uniform gossip algorithm also solves the rendezvous problem. We then use a multi-agent simulation to conduct a series of simulation experiments to compare the performance between the buffered and uniform gossip algorithms. Our results suggest that the buffered gossip algorithm can solve the rendezvous problem faster than the uniform gossip algorithm; however, the relative performance between these two solutions depends on the specific constraints of the problem and the parameters of the buffered gossip algorithm. PMID:25397882

Angles-only navigation for autonomous orbital rendezvous

NASA Astrophysics Data System (ADS)

Woffinden, David C.

The proposed thesis of this dissertation has both a practical element and theoretical component which aim to answer key questions related to the use of angles-only navigation for autonomous orbital rendezvous. The first and fundamental principle to this work argues that an angles-only navigation filter can determine the relative position and orientation (pose) between two spacecraft to perform the necessary maneuvers and close proximity operations for autonomous orbital rendezvous. Second, the implementation of angles-only navigation for on-orbit applications is looked upon with skeptical eyes because of its perceived limitation of determining the relative range between two vehicles. This assumed, yet little understood subtlety can be formally characterized with a closed-form analytical observability criteria which specifies the necessary and sufficient conditions for determining the relative position and velocity with only angular measurements. With a mathematical expression of the observability criteria, it can be used to (1) identify the orbital rendezvous trajectories and maneuvers that ensure the relative position and velocity are observable for angles-only navigation, (2) quantify the degree or level of observability and (3) compute optimal maneuvers that maximize observability. In summary, the objective of this dissertation is to provide both a practical and theoretical foundation for the advancement of autonomous orbital rendezvous through the use of angles-only navigation.

Starshade Rendezvous Mission Probe Concept

NASA Astrophysics Data System (ADS)

Seager, Sara; Kasdin, Jeremy; Starshade Rendezvous Probe Team

2018-01-01

The Starshade Rendezvous Mission Concept Prove is a Starshade that works with the WFIRST Mission, but is built and launched separately, with a rendezvous on orbit. A 2015 Exo-S report first detailed the mission concept. In the current study we develop a new scientific vision for WFIRST exoplanet discovery and characterization, using the complementary coronagraph and starshade to execute the most sensitive and thorough direct imaging campaign ever attempted. The overarching goal of our proposal is to carry out the first “deep dive” direct imaging exploration of planetary systems orbiting the nearest sun-like stars in a search for Earth-like planets using only a fraction of the WFIRST telescope time. The study aims to improve on the Exo-S 2015 report with updated study of the key spacecraft and starshade technology development issues, as related to WFIRST design changes since 2015 that make the timely implementation of such a mission possible.

Integrated Docking Simulation and Testing with the Johnson Space Center Six-Degree of Freedom Dynamic Test System

NASA Technical Reports Server (NTRS)

Mitchell, Jennifer D.; Cryan, Scott P.; Baker, Kenneth; Martin, Toby; Goode, Robert; Key, Kevin W.; Manning, Thomas; Chien, Chiun-Hong

2008-01-01

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as Automated Rendezvous and Docking, AR&D). The crewed versions may also perform AR&D, possibly with a different level of automation and/or autonomy, and must also provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success of the Constellation Program; this is carried as one of the CEV Project top risks. The Exploration Technology Development Program (ETDP) AR&D Sensor Technology Project seeks to reduce this risk by increasing technology maturation of selected relative navigation sensor technologies through testing and simulation. One of the project activities is a series of "pathfinder" testing and simulation activities to integrate relative navigation sensors with the Johnson Space Center Six-Degree-of-Freedom Test System (SDTS). The SDTS will be the primary testing location for the Orion spacecraft s Low Impact Docking System (LIDS). Project team members have integrated the Orion simulation with the SDTS computer system so that real-time closed loop testing can be performed with relative navigation sensors and the docking system in the loop during docking and undocking scenarios. Two relative navigation sensors are being used as part of a "pathfinder" activity in order to pave the way for future testing with the actual Orion sensors. This paper describes the test configuration and test results.

SERT D spacecraft study. [project planning and objectives

NASA Technical Reports Server (NTRS)

1974-01-01

The SERT D (Space Electric Rocket Test - D) study defines a possible spacecraft project that would demonstrate the use of electric ion thrusters for long-term (5 yr) station keeping and attitude control of a synchronous orbit satellite. Other mission objectives included in the study were: station walking to satellite rendezvous and inspection, use of low cost attitude sensing system, use of an advanced solar array orientation and slip ring system, and an ion thruster integrated directly with a solar array power source. The SERT D spacecraft, if launched, will become SERT 3 the third space electric thruster test.

Antares Post Launch Press Conference

NASA Image and Video Library

2013-09-18

Alan Lindenmoyer, program manager, NASA's Commercial Crew and Cargo Program, talks during a press conference held after the successful launch of the Orbital Sciences Corporation Antares rocket, with the Cygnus cargo spacecraft aboard, Wednesday, Sept. 18, 2013, NASA Wallops Flight Facility, Virginia. Cygnus is on its way to rendezvous with the space station. The spacecraft will deliver about 1,300 pounds (589 kilograms) of cargo, including food and clothing, to the Expedition 37 crew. Photo Credit: (NASA/Bill Ingalls)

A New Guidance Method for a Delta V and Re-entry Constrained Orbit Transfer Problem

DTIC Science & Technology

2005-06-01

a vehicle that undertakes a maneuver with the objective of precisely flying through a point in space at a particular time. The spacecraft must...for the Example Spacecraft . . . . 50 4-1 Graphical Results of Large Changes in Orbital Velocity . . . . . . . . . . . 62 4-2 Contours of Perigee...Orbit Relative to Rendezvous Point . . . . . . . . . . 98 6-2 Angular Rate and Angles for GEM-CR Maneuver with ∆θ = 90◦ . . . . . . 101 6-3 Position

KSC-2012-1851

NASA Image and Video Library

2012-02-17

Project Gemini: On Jan. 3, 1962, NASA announced the advanced Mercury Mark II project had been named "Gemini." After 12 missions – 2 uncrewed and 10 crewed – Project Gemini ended Nov. 15, 1966, following a nearly four-day, 59 orbit-flight. Its achievements included long-duration spaceflight, rendezvous and docking of two spacecraft in Earth orbit, extravehicular activity, and precision-controlled re-entry and landing of the spacecraft. Poster designed by Kennedy Space Center Graphics Department/Greg Lee. Credit: NASA

Analyses of the P/Wild 2 Images from STARDUST

NASA Technical Reports Server (NTRS)

Duxbury, Thomas C.

2004-01-01

This viewgraph presentation reviews the design of the Stardust spacecraft, and the trajectory that took it to rendezvous with the comet, Wild-2. Included are views of the comet, and comparisons with other astronomical bodies. Close up views show size, shape and orientation.

Launch of Agena Target Docking Vehicle atop Atlas launch vehicle

NASA Technical Reports Server (NTRS)

1966-01-01

An Agena Target Docking Vehicle atop its Atlas launch vehicle was launched fromt the Kennedy Space Center's Launch Complex 14 at 6:05 a.m., September 12, 1966. The Agena served as a rendezvous and docking vehicle for the Gemini 11 spacecraft.

Robust adaptive backstepping neural networks control for spacecraft rendezvous and docking with input saturation.

PubMed

Xia, Kewei; Huo, Wei

2016-05-01

This paper presents a robust adaptive neural networks control strategy for spacecraft rendezvous and docking with the coupled position and attitude dynamics under input saturation. Backstepping technique is applied to design a relative attitude controller and a relative position controller, respectively. The dynamics uncertainties are approximated by radial basis function neural networks (RBFNNs). A novel switching controller consists of an adaptive neural networks controller dominating in its active region combined with an extra robust controller to avoid invalidation of the RBFNNs destroying stability of the system outside the neural active region. An auxiliary signal is introduced to compensate the input saturation with anti-windup technique, and a command filter is employed to approximate derivative of the virtual control in the backstepping procedure. Globally uniformly ultimately bounded of the relative states is proved via Lyapunov theory. Simulation example demonstrates effectiveness of the proposed control scheme. Copyright © 2016 ISA. Published by Elsevier Ltd. All rights reserved.

Autonomous rendezvous and docking: A commercial approach to on-orbit technology validation

NASA Technical Reports Server (NTRS)

Tchoryk, Peter, Jr.; Dobbs, Michael E.; Conrad, David J.; Apley, Dale J.; Whitten, Raymond P.

1991-01-01

The Space Automation and Robotics Center (SpARC), a NASA-sponsored Center for the Commercial Development of Space (CCDS), in conjunction with its corporate affiliates, is planning an on-orbit validation of autonomous rendezvous and docking (ARD) technology. The emphasis in this program is to utilize existing technology and commercially available components whenever possible. The primary subsystems that will be validated by this demonstration include GPS receivers for navigation, a video-based sensor for proximity operations, a fluid connector mechanism to demonstrate fluid resupply capability, and a compliant, single-point docking mechanism. The focus for this initial experiment will be expendable launch vehicle (ELV) based and will make use of two residual Commercial Experiment Transporter (COMET) service modules. The first COMET spacecraft will be launched in late 1992 and will serve as the target vehicle. The ARD demonstration will take place in late 1994, after the second COMET spacecraft has been launched. The service module from the second COMET will serve as the chase vehicle.

Solar Sail Application to Comet Nucleus Sample Return

NASA Technical Reports Server (NTRS)

Taylor, Travis S.; Moton, Tryshanda T.; Robinson, Don; Anding, R. Charles; Matloff, Gregory L.; Garbe, Gregory; Montgomery, Edward

2003-01-01

Many comets have perihelions at distances within 1.0 Astronomical Unit (AU) from the sun. These comets typically are inclined out of the ecliptic. We propose that a solar sail spacecraft could be used to increase the inclination of the orbit to match that of these 1.0 AU comets. The solar sail spacecraft would match the orbit velocity for a short period of time, which would be long enough for a container to be injected into the comet's nucleus. The container would be extended from a long durable tether so that the solar sail would not be required to enter into the potentially degrading environment of the comet s atmosphere. Once the container has been filled with sample material, the tether is retracted. The solar sail would then lower its inclination and fly back to Earth for the sample return. In this paper, we describe the selection of cometary targets, the mission design, and the solar sailcraft design suitable for sail-comet rendezvous as well as possible rendezvous scenarios.

Advanced Video Guidance Sensor and next-generation autonomous docking sensors

NASA Astrophysics Data System (ADS)

Granade, Stephen R.

2004-09-01

In recent decades, NASA's interest in spacecraft rendezvous and proximity operations has grown. Additional instrumentation is needed to improve manned docking operations' safety, as well as to enable telerobotic operation of spacecraft or completely autonomous rendezvous and docking. To address this need, Advanced Optical Systems, Inc., Orbital Sciences Corporation, and Marshall Space Flight Center have developed the Advanced Video Guidance Sensor (AVGS) under the auspices of the Demonstration of Autonomous Rendezvous Technology (DART) program. Given a cooperative target comprising several retro-reflectors, AVGS provides six-degree-of-freedom information at ranges of up to 300 meters for the DART target. It does so by imaging the target, then performing pattern recognition on the resulting image. Longer range operation is possible through different target geometries. Now that AVGS is being readied for its test flight in 2004, the question is: what next? Modifications can be made to AVGS, including different pattern recognition algorithms and changes to the retro-reflector targets, to make it more robust and accurate. AVGS could be coupled with other space-qualified sensors, such as a laser range-and-bearing finder, that would operate at longer ranges. Different target configurations, including the use of active targets, could result in significant miniaturization over the current AVGS package. We will discuss these and other possibilities for a next-generation docking sensor or sensor suite that involve AVGS.

Advanced Video Guidance Sensor and Next Generation Autonomous Docking Sensors

NASA Technical Reports Server (NTRS)

Granade, Stephen R.

2004-01-01

In recent decades, NASA's interest in spacecraft rendezvous and proximity operations has grown. Additional instrumentation is needed to improve manned docking operations' safety, as well as to enable telerobotic operation of spacecraft or completely autonomous rendezvous and docking. To address this need, Advanced Optical Systems, Inc., Orbital Sciences Corporation, and Marshall Space Flight Center have developed the Advanced Video Guidance Sensor (AVGS) under the auspices of the Demonstration of Autonomous Rendezvous Technology (DART) program. Given a cooperative target comprising several retro-reflectors, AVGS provides six-degree-of-freedom information at ranges of up to 300 meters for the DART target. It does so by imaging the target, then performing pattern recognition on the resulting image. Longer range operation is possible through different target geometries. Now that AVGS is being readied for its test flight in 2004, the question is: what next? Modifications can be made to AVGS, including different pattern recognition algorithms and changes to the retro-reflector targets, to make it more robust and accurate. AVGS could be coupled with other space-qualified sensors, such as a laser range-and-bearing finder, that would operate at longer ranges. Different target configurations, including the use of active targets, could result in significant miniaturization over the current AVGS package. We will discuss these and other possibilities for a next-generation docking sensor or sensor suite that involve AVGS.

Unmanned planetary spacecraft chemical rocket propulsion.

NASA Technical Reports Server (NTRS)

Burlage, H., Jr.; Gin, W.; Riebling, R. W.

1972-01-01

Review of some chemical propulsion technology advances suitable for future unmanned spacecraft applications. Discussed system varieties include liquid space-storable propulsion systems, advanced liquid monopropellant systems, liquid systems for rendezvous and landing applications, and low-thrust high-performance solid-propellant systems, as well as hybrid space-storable systems. To optimize the performance and operational characteristics of an unmanned interplanetary spacecraft for a particular mission, and to achieve high cost effectiveness of the entire system, it is shown to be essential that the type of spacecraft propulsion system to be used matches, as closely as possible the various requirements and constraints. The systems discussed are deemed to be the most promising candidates for some of the anticipated interplanetary missions.

Apollo-Soyuz pamphlet no. 1: The flight

NASA Technical Reports Server (NTRS)

Page, L. W.; From, T. P.

1977-01-01

The goals of the Apollo-Soyuz Test Project are described in this first in a series of nine pamphlets designed as a curriculum supplement for teachers, supervisors, curriculum specialists, and textbook writers as well as for the general public. Aspects of the space flight covered include descriptions of the astronaut-cosmonaut meeting and of the spacecraft and landing module; spacecraft launch; control, and rendezvous; crew work schedule; and telemetry. Experiments performed are listed in tables, and their major results are summarized.

MISSION CONTROL CENTER (MCC) - APOLLO-SOYUZ TEST PROJECT (ASTP) - JSC

NASA Image and Video Library

1975-07-17

S75-28683 (17 July 1975) --- An overall view of the Mission Operations Control Room in the Mission Control Center during the joint U.S.-USSR Apollo-Soyuz Test Project docking mission in Earth orbit. M.P. Frank, the American senior ASTP flight director, is seated at his console in the right foreground. He is watching the large television monitor which shows a view of the Soyuz spacecraft as seen from the Apollo spacecraft during rendezvous and docking maneuvers.

Solar electric propulsion thruster interactions with solar arrays

NASA Technical Reports Server (NTRS)

Parks, D. E.; Katz, I.

1977-01-01

The effect of interactions of spacecraft-generated and naturally occurring plasmas with high voltage solar array components on an advanced solar electric propulsion system proposed for the Halley's Comet rendezvous mission was investigated. The spacecraft-generated plasma consists of mercury ions and neutralizing electrons resulting from the operation of ion thrusters (the charge-exchange plasma) and associated hollow cathode neutralizers. Quantitative results are given for the parasitic currents and power coupled into solar arrays with voltage fixed as a function of position on the array.

Antares Post Launch Press Conference

NASA Image and Video Library

2013-09-18

Alan Lindenmoyer, program manager, NASA's Commercial Crew and Cargo Program, left, and, Frank Culbertson, executive vice president, Orbital Sciences Corporation,are seen during a press conference held after the successful launch of the Orbital Sciences Antares rocket, with the Cygnus cargo spacecraft aboard, Wednesday, Sept. 18, 2013, NASA Wallops Flight Facility, Virginia. Cygnus is on its way to rendezvous with the space station. The spacecraft will deliver about 1,300 pounds (589 kilograms) of cargo, including food and clothing, to the Expedition 37 crew. Photo Credit: (NASA/Bill Ingalls)

Cost efficient operations for Discovery class missions

NASA Technical Reports Server (NTRS)

Cameron, G. E.; Landshof, J. A.; Whitworth, G. W.

1994-01-01

The Near Earth Asteroid Rendezvous (NEAR) program at The Johns Hopkins University Applied Physics Laboratory is scheduled to launch the first spacecraft in NASA's Discovery program. The Discovery program is to promote low cost spacecraft design, development, and mission operations for planetary space missions. The authors describe the NEAR mission and discuss the design and development of the NEAR Mission Operations System and the NEAR Ground System with an emphasis on those aspects of the design that are conducive to low-cost operations.

Apollo Soyuz test project, USA-USSR. [mission plan of spacecraft docking

NASA Technical Reports Server (NTRS)

1975-01-01

The mission plan of the docking of a United States Apollo and a Soviet Union Soyuz spacecraft in Earth orbit to test compatible rendezvous and docking equipment and procedures is presented. Space experiments conducted jointly by the astronauts and cosmonauts during the joint phase of the mission as well as experiments performed solely by the U.S. astronauts and spread over the nine day span of the flight are included. Biographies of the astronauts and cosmonauts are given.

The Cluster Orbits With Perturbations of Keplerian Elements (COWPOKE) Equations

DTIC Science & Technology

2004-03-20

Hill’s equations2 (also known as the Clohessy - Wiltshire equations3). Hill’s equations describe the relative motion of spacecraft using a spacecraft- 2...January 2002, AAS 02-115. 2. Hill, G. W., “Researches in the Lunar Theory,” American Journal of Mathematics, Vol. 1, No. 1, 1878, pp. 5-26. 3. Clohessy ...W. H., R. S. Wiltshire , “Terminal Guidance System for Satellite Rendezvous,” Journal of the Aerospace Sciences, September 1960, pp. 653-658 4. Gim

Main-belt asteroid exploration - Mission options for the 1990s

NASA Technical Reports Server (NTRS)

Yen, Chen-Wan L.

1989-01-01

An extensive investigation of the ways to rendezvous with diverse groups of asteroids residing between 2.0 and 5.0 AU is made, and the extent of achievable missions using the STS upper-stage launch vehicles (IUS 2-Stage/Star-48 or NASA Centaur) is examined. With judicious use of earth, Mars, and Jupiter gravity assists, rendezvous with some asteroids in all regions of space is possible. It is also shown that the STS upper stages are capable of carrying out missions beyond a single rendezvous, namely with several flybys and/or multiple rendezvous.

Ground Simulation of an Autonomous Satellite Rendezvous and Tracking System Using Dual Robotic Systems

NASA Technical Reports Server (NTRS)

Trube, Matthew J.; Hyslop, Andrew M.; Carignan, Craig R.; Easley, Joseph W.

2012-01-01

A hardware-in-the-loop ground system was developed for simulating a robotic servicer spacecraft tracking a target satellite at short range. A relative navigation sensor package "Argon" is mounted on the end-effector of a Fanuc 430 manipulator, which functions as the base platform of the robotic spacecraft servicer. Machine vision algorithms estimate the pose of the target spacecraft, mounted on a Rotopod R-2000 platform, relay the solution to a simulation of the servicer spacecraft running in "Freespace", which performs guidance, navigation and control functions, integrates dynamics, and issues motion commands to a Fanuc platform controller so that it tracks the simulated servicer spacecraft. Results will be reviewed for several satellite motion scenarios at different ranges. Key words: robotics, satellite, servicing, guidance, navigation, tracking, control, docking.

Report of the Attitude Control and Attitude Determination Panel. [spacecraft instrumentation technology

NASA Technical Reports Server (NTRS)

1979-01-01

Failures and deficiencies in flight programs are reviewed and suggestions are made for avoiding them. The technology development problem areas considered are control configured vehicle design, gyros, solid state star sensors, control instrumentation, tolerant/accomodating control systems, large momentum exchange devices, and autonomous rendezvous and docking.

ISS during STS-135 Approach

NASA Image and Video Library

2011-07-10

S135-E-006777 (10 July 2011) --- This is one of a series of images showing the International Space Station photographed by a crewmember onboard the space shuttle Atlantis as the two spacecraft performed rendezvous and docking operations on the STS-135 mission's third day in Earth orbit. Photo credit: NASA

ISS during STS-135 Approach

NASA Image and Video Library

2011-07-10

S135-E-006784 (10 July 2011) --- This is one of a series of images showing the International Space Station photographed by a crewmember onboard the space shuttle Atlantis as the two spacecraft performed rendezvous and docking operations on the STS-135 mission's third day in Earth orbit. Photo credit: NASA

ISS Segments during STS-135 Approach

NASA Image and Video Library

2011-07-10

S135-E-006787 (10 July 2011) --- This is one of a series of images showing the International Space Station photographed by a crewmember onboard the space shuttle Atlantis as the two spacecraft performed rendezvous and docking operations on the STS-135 mission's third day in Earth orbit. Photo credit: NASA

ISS during STS-135 Approach

NASA Image and Video Library

2011-07-10

S135-E-006700 (10 July 2011) --- This is one of a series of images showing the International Space Station photographed by a crewmember onboard the space shuttle Atlantis as the two spacecraft performed rendezvous and docking operations on the STS-135 mission's third day in Earth orbit. Photo credit: NASA

ISS during STS-135 Approach

NASA Image and Video Library

2011-07-10

S135-E-006698 (10 July 2011) --- This is one of a series of images showing the International Space Station photographed by a crewmember onboard the space shuttle Atlantis as the two spacecraft performed rendezvous and docking operations on the STS-135 mission's third day in Earth orbit. Photo credit: NASA

ISS during STS-135 Approach

NASA Image and Video Library

2011-07-10

S135-E-006702 (10 July 2011) --- This is one of a series of images showing the International Space Station photographed by a crewmember onboard the space shuttle Atlantis as the two spacecraft performed rendezvous and docking operations on the STS-135 mission's third day in Earth orbit. Photo credit: NASA

Requirements and Capabilities for Planetary Missions: Mariner Encke Ballistic Flyby 1980

NASA Technical Reports Server (NTRS)

Ball, G. G.; Bird, T. H.

1975-01-01

This mission will provide a broad-based fast reconnaissance of comet Encke, building a data base for subsequent more detailed comet investigations, including rendezvous. After a 3 month flight, the spacecraft will encounter the comet at a nominal range of about 500 km. Flyby velocity will be 7 to 28 km/sec depending on choice of arrival data (0 to 35 days before Encke perihelion) and launch vehicle. The spacecraft will be similar to the MVM 73 spacecraft, with scan platform and 117 kbps encounter data rate, and designed to survive the thermal environment of 0.34 to 0.8 AU.

Apollo 7/S-IVB Rendezvous in space

NASA Technical Reports Server (NTRS)

1968-01-01

The expended Saturn IVB stage as photographed from the Apollo 7 spacecraft during transposition and docking maneuvers at an altitude of 126 nautical miles, at ground elapsed time of three hours, 11 minutes. The round, white disc inside the open panels of the Saturn IVB is a simulated docking target similar to that used on the lunar module for docking during lunar missions. The spacecraft is directly over Odessa-Midland, Texas. The view between the two panels (area of large puffy clouds) extends southwest across Texas into the Mexican State of Chihuahua. The distance between the Apollo 7 spacecraft and the S-(VB is approximately 50 feet.

Interplanetary spacecraft design using solar electric propulsion

NASA Technical Reports Server (NTRS)

Duxbury, J. H.; Paul, G. M.

1974-01-01

Emphasis of the electric propulsion technology program is now on the application of solar electric propulsion to scientific missions. Candidate planetary, cometary, and geosynchronous missions are being studied. The object of this paper is to describe a basic spacecraft design proposed as the means to accomplish (1) a comet Encke slow flyby, (2) a comet Encke rendezvous, and (3) an out-of-the-ecliptic mission. The discussion includes design differences foreseen for the various missions and indicates those areas where spacecraft design commonality is possible. Particular emphasis is placed on a solar electric propulsion module design which permits an attractive degree of design inheritance from mission to mission.

Low-energy multiple rendezvous of main belt asteroids

NASA Technical Reports Server (NTRS)

Penzo, Paul A.; Bender, David F.

1992-01-01

An approach to multiple asteroid rendezvous missions to the main belt region is proposed. In this approach key information which consists of a launch date and delta V can be generated for all possible pairs of asteroids satisfying specific constraints. This information is made available on a computer file for 1000 numbered asteroids with reasonable assumptions, limitations, and approximations to limit the computer requirements and the size of the data file.

Gemini Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1962-09-21

A researcher at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a small-scale model of the Gemini capsule in the 10- by 10-Foot Supersonic Wind Tunnel test section. Gemini was added to NASA’s manned space program after its predecessor, Mercury, and its antecedent, Apollo, were already established. Gemini was a transitional mission designed provide the astronauts with practice docking with other spacecraft and withstanding durations in space up to two weeks. The program was officially announced on December 7, 1961, but planning began in mid-1959. It was named Gemini after the zodiac twins because of the spacecraft’s two passenger capacity. The Gemini Program was the first program to start at the new Manned Spacecraft Center in Houston, now the Johnson Space Center. Unlike Mercury and Apollo, Lewis had very little involvement with the Gemini Program. This model was tested in the 10- by 10 tunnel for several weeks in September 1962. Lewis began managing the Agena second-stage rocket program shortly after this photograph was taken. Agenas were used to launch a variety of spacecraft and satellites in the 1960s. They were also used on several Gemini missions to provide targets for the astronauts to practice their rendezvous maneuvers. Gemini had two unmanned and ten manned flights in 1965 and 1966. These yielded the first spacewalks, long-duration space missions, first onboard computer, docking with a second spacecraft, and rendezvous maneuvers.

Mars Sample Return - Launch and Detection Strategies for Orbital Rendezvous

NASA Technical Reports Server (NTRS)

Woolley, Ryan C.; Mattingly, Richard L.; Riedel, Joseph E.; Sturm, Erick J.

2011-01-01

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/cache rover in 2018, an orbiter with an Earth return vehicle in 2022, and a fetch rover and ascent vehicle in 2024. Strategies are presented to launch the sample into a coplanar orbit with the Orbiter which facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits exist at 457 and 572 km which provide multiple launch opportunities with similar geometries for detection and rendezvous.

Mars Sample Return: Launch and Detection Strategies for Orbital Rendezvous

NASA Technical Reports Server (NTRS)

Woolley, Ryan C.; Mattingly, Richard L.; Riedel, Joseph E.; Sturm, Erick J.

2011-01-01

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/ caching rover in 2018, an Earth return orbiter in 2022, and a fetch rover with ascent vehicle in 2024. Strategies are presented to launch the sample into a nearly coplanar orbit with the Orbiter which would facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits existat 457 and 572 km which would provide multiple launch opportunities with similar geometries for detection and rendezvous.

Development of a miniature scanning electron microscope for in-flight analysis of comet dust

NASA Technical Reports Server (NTRS)

Conley, J. M.; Bradley, J. G.; Giffin, C. E.; Albee, A. L.; Tomassian, A. D.

1983-01-01

A description is presented of an instrument which was developed with the original goal of being flown on the International Comet Mission, scheduled for a 1985 launch. The Scanning Electron Microscope and Particle Analyzer (SEMPA) electron miniprobe is a miniaturized electrostatically focused electron microscope and energy dispersive X-ray analyzer for in-flight analysis of comet dust particles. It was designed to be flown on board a comet rendezvous spacecraft. Other potential applications are related to asteroid rendezvous and planetary lander missions. According to the development objectives, SEMPA miniprobe is to have the capability for imaging and elemental analysis of particles in the size range of 0.25 microns and larger.

Report of the Task Force on the Shuttle-Mir Rendezvous and Docking Missions

NASA Technical Reports Server (NTRS)

1994-01-01

In October 1992, Russia and the U.S. agreed to conduct a fundamentally new program of human cooperation in space. This original 'Shuttle-Mir' project encompassed combined astronaut-cosmonaut activities on the Shuttle, Soyuz, and Mir spacecraft. At that time, the project was limited to: the STS-60 Shuttle mission, which was completed in February 1994 and carried the first Russian cosmonaut; the planned March 1995 Soyuz 18 launch which will carry a U.S. astronaut to the Mir space station for a three month mission; and the STS-71 Shuttle mission which is scheduled to rendezvous and dock with the Mir space station in June 1995. The Task Force's specific recommendations are given.

Selected tether applications in space: An analysis of five selected concepts

NASA Technical Reports Server (NTRS)

1984-01-01

Ground rules and assumptions; operations; orbit considerations/dynamics; tether system design and dynamics; functional requirements; hardware concepts; and safety factors are examined for five scenarios: tethered effected separation of an Earth bound shuttle from the space station; tether effected orbit boost of a spacecraft (AXAF) into its operational orbit from the shuttle; an operational science/technology platform tether deployed from space station; a tether mediated rendezvous involving an OMV tether deployed from space station to rendezvous with an aerobraked OTV returning to geosynchronous orbit from a payload delivery mission; and an electrodynamic tether used in a dual motor/generator mode to serve as the primary energy storage facility for space station.

KU-Band rendezvous radar performance computer simulation model

NASA Technical Reports Server (NTRS)

Griffin, J. W.

1980-01-01

The preparation of a real time computer simulation model of the KU band rendezvous radar to be integrated into the shuttle mission simulator (SMS), the shuttle engineering simulator (SES), and the shuttle avionics integration laboratory (SAIL) simulator is described. To meet crew training requirements a radar tracking performance model, and a target modeling method were developed. The parent simulation/radar simulation interface requirements, and the method selected to model target scattering properties, including an application of this method to the SPAS spacecraft are described. The radar search and acquisition mode performance model and the radar track mode signal processor model are examined and analyzed. The angle, angle rate, range, and range rate tracking loops are also discussed.

Gemini Program Mission Report for Gemini-Titan 1 (GT-1)

NASA Technical Reports Server (NTRS)

1964-01-01

The Gemini-Titan 1 (GT-1) space vehicle was comprised of the Gemini spacecraft and the Gemini launch vehicle. The Gemini launch vehicle is a two-stage modified Titan II ICBM. The major modifications are the addition of a malfunction detection system and a secondary flight controls system. The Gemini spacecraft, designed to carry a crew of two men on earth orbital and rendezvous missions, was unmanned for the flight reported herein (GT-1). There were no complete Gemini flight systems on board; however, the C-band transponder and telemetry transmitters were Gemini flight subsystems. Dummy equipment, having a mass and moment of inertia equal to flight system equipment, was installed in the spacecraft. The Spacecraft was instrumented to obtain data on spacecraft heating, structural loading, vibration, sound pressure levels, and temperature and pressure during the launch phase.

Ku-Band rendezvous radar performance computer simulation model

NASA Technical Reports Server (NTRS)

Magnusson, H. G.; Goff, M. F.

1984-01-01

All work performed on the Ku-band rendezvous radar performance computer simulation model program since the release of the preliminary final report is summarized. Developments on the program fall into three distinct categories: (1) modifications to the existing Ku-band radar tracking performance computer model; (2) the addition of a highly accurate, nonrealtime search and acquisition performance computer model to the total software package developed on this program; and (3) development of radar cross section (RCS) computation models for three additional satellites. All changes in the tracking model involved improvements in the automatic gain control (AGC) and the radar signal strength (RSS) computer models. Although the search and acquisition computer models were developed under the auspices of the Hughes Aircraft Company Ku-Band Integrated Radar and Communications Subsystem program office, they have been supplied to NASA as part of the Ku-band radar performance comuter model package. Their purpose is to predict Ku-band acquisition performance for specific satellite targets on specific missions. The RCS models were developed for three satellites: the Long Duration Exposure Facility (LDEF) spacecraft, the Solar Maximum Mission (SMM) spacecraft, and the Space Telescopes.

Ku-Band rendezvous radar performance computer simulation model

NASA Astrophysics Data System (ADS)

Magnusson, H. G.; Goff, M. F.

1984-06-01

All work performed on the Ku-band rendezvous radar performance computer simulation model program since the release of the preliminary final report is summarized. Developments on the program fall into three distinct categories: (1) modifications to the existing Ku-band radar tracking performance computer model; (2) the addition of a highly accurate, nonrealtime search and acquisition performance computer model to the total software package developed on this program; and (3) development of radar cross section (RCS) computation models for three additional satellites. All changes in the tracking model involved improvements in the automatic gain control (AGC) and the radar signal strength (RSS) computer models. Although the search and acquisition computer models were developed under the auspices of the Hughes Aircraft Company Ku-Band Integrated Radar and Communications Subsystem program office, they have been supplied to NASA as part of the Ku-band radar performance comuter model package. Their purpose is to predict Ku-band acquisition performance for specific satellite targets on specific missions. The RCS models were developed for three satellites: the Long Duration Exposure Facility (LDEF) spacecraft, the Solar Maximum Mission (SMM) spacecraft, and the Space Telescopes.

GEMINI-TITAN (GT)-9 PREFLIGHT ACTIVITY - ASTRONAUT THOMAS P. STAFFORD - MISC. - KSC

NASA Image and Video Library

1969-01-21

S66-32044 (17 May 1966) --- Astronauts Eugene A. Cernan (left), pilot, and Thomas P. Stafford, command pilot, discuss the postponed Gemini-9 mission just after egressing their spacecraft in the white room atop Pad 19. The Agena Target Vehicle failed to achieve orbit, causing a termination of the mission. The spaceflight (to be called Gemini-9A) has been rescheduled for May 31. A Gemini Augmented Target Docking Adapter will be used as the rendezvous and docking vehicle for the Gemini-9 spacecraft. Photo credit: NASA

MISSION CONTROL CENTER (MCC) - APOLLO-SOYUZ TEST PROJECT (ASTP) - JSC

NASA Image and Video Library

1975-07-17

S75-28682 (17 July 1975) --- An overall view of the Mission Operations Control Room in the Mission Control Center during the joint U.S.-USSR Apollo-Soyuz Test Project docking mission in Earth orbit. The large television monitor shows a view of the Soyuz spacecraft as seen from the Apollo spacecraft during rendezvous and docking maneuvers. Eugene F. Kranz, JSC Deputy Director of Flight Operations, is standing in the foreground. M.P. Frank, the American senior ASTP flight director, is partially obscured on the right.

Apollo 7/S-IVB Rendezvous in space

NASA Image and Video Library

1968-10-11

AS07-03-1535 (11 Oct. 1968) --- The expended Saturn IVB stage as photographed from the Apollo 7 spacecraft during transposition and docking maneuvers at an altitude of 126 nautical miles, at ground elapsed time of three hours, 11 minutes. The round, white disc inside the open panels of the Saturn IVB is a simulated docking target similar to that used on the lunar module for docking during lunar missions. The spacecraft is directly over Odessa-Midland, Texas. The view between the two panels (area of large puffy clouds) extends southwest across Texas into the Mexican State of Chihuahua. The distance between the Apollo 7 spacecraft and the S-IVB is approximately 50 feet.

SEL2 servicing: increased science return via on-orbit propellant replenishment

NASA Astrophysics Data System (ADS)

Reed, Benjamin B.; DeWeese, Keith; Kienlen, Michael; Aranyos, Thomas; Pellegrino, Joseph; Bacon, Charles; Qureshi, Atif

2016-07-01

Spacecraft designers are driving observatories to the distant Sun-Earth Lagrange Point 2 (SEL2) to meet ever-increasing science requirements. The mass fraction dedicated to propellant for these observatories to reach and operate at SEL2 will be allocated with the upmost care, as it comes at the expense of optics and instrument masses. As such, these observatories could benefit from on-orbit refueling, allowing greater dry-to-wet mass ratio at launch and/or longer mission life. NASA is developing technologies, capabilities and integrated mission designs for multiple servicing applications in low Earth orbit (LEO), geosynchronous Earth orbit (GEO) and cisluner locations. Restore-L, a mission officially in formulation, will launch a free-flying robotic servicer to refuel a government-owned satellite in LEO by mid 2020. This paper will detail the results of a point design mission study to extend Restore-L servicing technologies from LEO to SEL2. This SEL2 mission would launch an autonomous, robotic servicer spacecraft equipped to extend the life of two space assets through refueling. Two space platforms were chosen to 1) drive the requirements for achieving SEL2 orbit and rendezvous with a spacecraft, and 2) to drive the requirements to translate within SEL2 to conduct a follow-on servicing mission. Two fuels, xenon and hydrazine, were selected to assess a multiple delivery system. This paper will address key mission drivers, such as servicer autonomy (necessitated due to communications latency at L2). Also discussed will be the value of adding cooperative servicing elements to the client observatories to reduce mission risk.

Enhanced Gravity Tractor Technique for Planetary Defense

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D.; Reeves, David M.; Hopkins, Joshua B.; Wade, Darren W.; Tantardini, Marco; Shen, Haijun

2015-01-01

Given sufficient warning time, Earth-impacting asteroids and comets can be deflected with a variety of different "slow push/pull" techniques. The gravity tractor is one technique that uses the gravitational attraction of a rendezvous spacecraft to the impactor and a low-thrust, high-efficiency propulsion system to provide a gradual velocity change and alter its trajectory. An innovation to this technique, known as the Enhanced Gravity Tractor (EGT), uses mass collected in-situ to augment the mass of the spacecraft, thereby greatly increasing the gravitational force between the objects. The collected material can be a single boulder, multiple boulders, regolith or a combination of different sources. The collected mass would likely range from tens to hundreds of metric tons depending on the size of the impactor and warning time available. Depending on the propulsion system's capability and the mass collected, the EGT approach can reduce the deflection times by a factor of 10 to 50 or more, thus reducing the deflection times of several decades to years or less and overcoming the main criticism of the traditional gravity tractor approach. Additionally, multiple spacecraft can orbit the target in formation to provide the necessary velocity change and further reduce the time needed by the EGT technique to divert hazardous asteroids and comets. The robotic segment of NASA's Asteroid Redirect Mission (ARM) will collect a multi-ton boulder from the surface of a large Near-Earth Asteroid (NEA) and will provide the first ever demonstration of the EGT technique and validate one method of collecting in-situ mass on an asteroid of hazardous size.

Avionic Architecture for Model Predictive Control Application in Mars Sample & Return Rendezvous Scenario

NASA Astrophysics Data System (ADS)

Saponara, M.; Tramutola, A.; Creten, P.; Hardy, J.; Philippe, C.

2013-08-01

Optimization-based control techniques such as Model Predictive Control (MPC) are considered extremely attractive for space rendezvous, proximity operations and capture applications that require high level of autonomy, optimal path planning and dynamic safety margins. Such control techniques require high-performance computational needs for solving large optimization problems. The development and implementation in a flight representative avionic architecture of a MPC based Guidance, Navigation and Control system has been investigated in the ESA R&T study “On-line Reconfiguration Control System and Avionics Architecture” (ORCSAT) of the Aurora programme. The paper presents the baseline HW and SW avionic architectures, and verification test results obtained with a customised RASTA spacecraft avionics development platform from Aeroflex Gaisler.

Mars, Phobos, and Deimos Sample Return Enabled by ARRM Alternative Trade Study Spacecraft

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew; Merrill, Raymond G.; Qu, Min; Naasz, Bo J.

2014-01-01

The Asteroid Robotic Redirect Mission (ARRM) has been the topic of many mission design studies since 2011. The reference ARRM spacecraft uses a powerful solar electric propulsion (SEP) system and a bag device to capture a small asteroid from an Earth-like orbit and redirect it to a distant retrograde orbit (DRO) around the moon. The ARRM Option B spacecraft uses the same propulsion system and multi-Degree of Freedom (DoF) manipulators device to retrieve a very large sample (thousands of kilograms) from a 100+ meter diameter farther-away Near Earth Asteroid (NEA). This study will demonstrate that the ARRM Option B spacecraft design can also be used to return samples from Mars and its moons - either by acquiring a large rock from the surface of Phobos or Deimos, and or by rendezvousing with a sample-return spacecraft launched from the surface of Mars.

Mars, Phobos, and Deimos Sample Return Enabled by ARRM Alternative Trade Study Spacecraft

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew; Naasz, Bo; Merill, Raymond G.; Qu, Min

2014-01-01

The Asteroid Robotic Redirect Mission (ARRM) has been the topic of many mission design studies since 2011. The reference ARRM spacecraft uses a powerful solar electric propulsion (SEP) system and a bag device to capture a small asteroid from an Earth-like orbit and redirect it to a distant retrograde orbit (DRO) around the moon. The ARRM Option B spacecraft uses the same propulsion system and multi-Degree of Freedom (DoF) manipulators device to retrieve a very large sample (thousands of kilograms) from a 100+ meter diameter farther-away Near Earth Asteroid (NEA). This study will demonstrate that the ARRM Option B spacecraft design can also be used to return samples from Mars and its moons - either by acquiring a large rock from the surface of Phobos or Deimos, and/or by rendezvousing with a sample-return spacecraft launched from the surface of Mars.

A Sampling Based Approach to Spacecraft Autonomous Maneuvering with Safety Specifications

NASA Technical Reports Server (NTRS)

Starek, Joseph A.; Barbee, Brent W.; Pavone, Marco

2015-01-01

This paper presents a methods for safe spacecraft autonomous maneuvering that leverages robotic motion-planning techniques to spacecraft control. Specifically the scenario we consider is an in-plan rendezvous of a chaser spacecraft in proximity to a target spacecraft at the origin of the Clohessy Wiltshire Hill frame. The trajectory for the chaser spacecraft is generated in a receding horizon fashion by executing a sampling based robotic motion planning algorithm name Fast Marching Trees (FMT) which efficiently grows a tree of trajectories over a set of probabillistically drawn samples in the state space. To enforce safety the tree is only grown over actively safe samples for which there exists a one-burn collision avoidance maneuver that circularizes the spacecraft orbit along a collision-free coasting arc and that can be executed under potential thrusters failures. The overall approach establishes a provably correct framework for the systematic encoding of safety specifications into the spacecraft trajectory generations process and appears amenable to real time implementation on orbit. Simulation results are presented for a two-fault tolerant spacecraft during autonomous approach to a single client in Low Earth Orbit.

Imaging Flash Lidar for Autonomous Safe Landing and Spacecraft Proximity Operation

NASA Technical Reports Server (NTRS)

Amzajerdian, Farzin; Roback, Vincent E.; Brewster, Paul F.; Hines, Glenn D.; Bulyshev, Alexander E.

2016-01-01

3-D Imaging flash lidar is recognized as a primary candidate sensor for safe precision landing on solar system bodies (Moon, Mars, Jupiter and Saturn moons, etc.), and autonomous rendezvous proximity operations and docking/capture necessary for asteroid sample return and redirect missions, spacecraft docking, satellite servicing, and space debris removal. During the final stages of landing, from about 1 km to 500 m above the ground, the flash lidar can generate 3-Dimensional images of the terrain to identify hazardous features such as craters, rocks, and steep slopes. The onboard fli1ght computer can then use the 3-D map of terrain to guide the vehicle to a safe location. As an automated rendezvous and docking sensor, the flash lidar can provide relative range, velocity, and bearing from an approaching spacecraft to another spacecraft or a space station from several kilometers distance. NASA Langley Research Center has developed and demonstrated a flash lidar sensor system capable of generating 16k pixels range images with 7 cm precision, at a 20 Hz frame rate, from a maximum slant range of 1800 m from the target area. This paper describes the lidar instrument design and capabilities as demonstrated by the closed-loop flight tests onboard a rocket-propelled free-flyer vehicle (Morpheus). Then a plan for continued advancement of the flash lidar technology will be explained. This proposed plan is aimed at the development of a common sensor that with a modest design adjustment can meet the needs of both landing and proximity operation and docking applications.

Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics

NASA Technical Reports Server (NTRS)

Burns, Richard D. (Inventor); Cepollina, Frank J. (Inventor); Jedhrich, Nicholas M. (Inventor); Holz, Jill M. (Inventor); Corbo, James E. (Inventor)

2008-01-01

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics

NASA Technical Reports Server (NTRS)

Burns, Richard D. (Inventor); Jedhrich, Nicholas M. (Inventor); Cepollina, Frank J. (Inventor); Holz, Jill M. (Inventor); Corbo, James E. (Inventor)

2007-01-01

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

NASA Technical Reports Server (NTRS)

Holz, Jill M. (Inventor); Corbo, James E. (Inventor); Burns, Richard D. (Inventor); Cepollina, Frank J. (Inventor); Jedhrich, Nicholas M. (Inventor)

2009-01-01

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

NASA Technical Reports Server (NTRS)

Burns, Richard D. (Inventor); Cepollina, Frank J. (Inventor); Jedhrich, Nicholas M. (Inventor); Holz, Jill M. (Inventor); Corbo, James E. (Inventor)

2007-01-01

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Agena Target Vehicle atop Atlas Launch vehicle launched from KSC

NASA Technical Reports Server (NTRS)

1966-01-01

An Agena Target Vehicle atop its Atlas Launch vehicle is launched from the Kennedy Space Center (KSC) Launch Complex 14 at 10:15 am.m., May 17, 1966. The Agena was intended as a rendezvous and docking vehicle for the Gemini 9 spacecraft. However, since the Agena failed to achieve orbit, the Gemini 9 mission was postponed.

The Exploration of Near-Earth Objects

NASA Astrophysics Data System (ADS)

1998-01-01

Near-Earth objects (NEOs) are asteroids and comets with orbits that intersect or pass near that of our planet. About 400 NEOs are currently known, but the entire population contains perhaps 3000 objects with diameters larger than 1 km. These objects, thought to be similar in many ways to the ancient planetesimal swarms that accreted to form the planets, are interesting and highly accessible targets for scientific research. They carry records of the solar system's birth and the geologic evolution of small bodies in the interplanetary region. Because collisions of NEOs with Earth pose a finite hazard to life, the exploration of these objects is particularly urgent. Devising appropriate risk-avoidance strategies requires quantitative characterization of NEOS. They may also serve as resources for use by future human exploration missions. The scientific goals of a focused NEO exploration program are to determine their orbital distribution, physical characteristics, composition, and origin. Physical characteristics, such as size, shape, and spin properties, have been measured for approximately 80 NEOs using observations at infrared, radar, and visible wavelengths. Mineralogical compositions of a comparable number of NEOs have been inferred from visible and near-infrared spectroscopy. The formation and geologic histories of NEOs and related main-belt asteroids are currently inferred from studies of meteorites and from Galileo and Near-Earth Asteroid Rendezvous spacecraft flybys of three main-belt asteroids. Some progress has also been made in associating specific types of meteorites with main-belt asteroids, which probably are the parent bodies of most NEOs. The levels of discovery of NEOs in the future will certainly increase because of the application of new detection systems. The rate of discovery may increase by an order of magnitude, allowing the majority of Earth-crossing asteroids and comets with diameters greater than 1 km to he discovered in the next decade. A small fraction of NEOs are particularly accessible for exploration by spacecraft. To identify the exploration targets of highest scientific interest, the orbits and classification of a large number of NEOs should be determined by telescopic observations. Desired characterization would also include measurements of size, mass, shape, surface composition and heterogeneity, gas and dust emission, and rotation. Laboratory studies of meteorites can focus NEO exploration objectives and quantify the information obtained from telescopes. Once high-priority targets have been identified, various kinds of spacecraft missions (flyby, rendezvous, and sample return) can be designed. Some currently operational (Near-Earth Asteroid Rendezvous [NEAR]) or planned (Deep Space 1) U.S. missions are of the first two types, and other planned U.S. and Japanese spacecraft missions will return samples. Rendezvous missions with sample return are particularly desirable from a scientific perspective because of the very great differences in the analytical capabilities that can be brought to bear in orbit and in the laboratory setting. Although it would be difficult to justify human exploration of NEOs on the basis of cost-benefit analysis of scientific results alone, a strong case can be made for starting with NEOs if the decision to carry out human exploration beyond low Earth orbit is made for other reasons. Some NEOs are especially attractive targets for astronaut missions because of their orbital accessibility and short flight duration. Because they represent deep space exploration at an intermediate level of technical challenge, these missions would also serve as stepping stones for human missions to Mars. Human exploration of NEOs would provide significant advances in observational and sampling capabilities. With respect to ground based telescopic studies, the recommended baseline is that NASA and other appropriate agencies suupport research programs for interpreting the spectra of near-Earth objects (NEOs), continue and coordinate currently supported surveys to discover and determine the orbits of NEOs and develop policies for the public disclosure of results relating to potential hazards. Augmentation to this baseline program include in priority order: (1) provide routine or priority access to existiing ground-based optical and infrared telescopes and radar facilities for characterization of NEOs during favorable encounters; or (2) provide expanded, dedicated telescope access for characterization of NEOs. Appropriate augmentations to existing programs include the following: (1) Develop technological advances in spacecraft capabilities, including nonchemical propulsion and autonomous navigation systems, low-power and low-mass anlaytical instrumentation for remote and in situ studies, and multiple penetrators and other sampling and sample-handling systems to allow low-cost rendezvous and sample return missions; and (2) study technical requirements for human expeditions to NEOs. Although studies evaluating the risk of asteroid collisions with Earth and the means of averting them are desirable, they are beyond the scope of this report.

Cost-Effective Icy Bodies Exploration using Small Satellite Missions

NASA Technical Reports Server (NTRS)

Jonsson, Jonas; Mauro, David; Stupl, Jan; Nayak, Michael; Aziz, Jonathan; Cohen, Aaron; Colaprete, Anthony; Dono-Perez, Andres; Frost, Chad; Klamm, Benjamin;

Comet nucleus and asteroid sample return missions

NASA Technical Reports Server (NTRS)

Melton, Robert G.; Thompson, Roger C.; Starchville, Thomas F., Jr.; Adams, C.; Aldo, A.; Dobson, K.; Flotta, C.; Gagliardino, J.; Lear, M.; Mcmillan, C.

1992-01-01

During the 1991-92 academic year, the Pennsylvania State University has developed three sample return missions: one to the nucleus of comet Wild 2, one to the asteroid Eros, and one to three asteroids located in the Main Belt. The primary objective of the comet nucleus sample return mission is to rendezvous with a short period comet and acquire a 10 kg sample for return to Earth. Upon rendezvous with the comet, a tethered coring and sampler drill will contact the surface and extract a two-meter core sample from the target site. Before the spacecraft returns to Earth, a monitoring penetrator containing scientific instruments will be deployed for gathering long-term data about the comet. A single asteroid sample return mission to the asteroid 433 Eros (chosen for proximity and launch opportunities) will extract a sample from the asteroid surface for return to Earth. To limit overall mission cost, most of the mission design uses current technologies, except the sampler drill design. The multiple asteroid sample return mission could best be characterized through its use of future technology including an optical communications system, a nuclear power reactor, and a low-thrust propulsion system. A low-thrust trajectory optimization code (QuickTop 2) obtained from the NASA LeRC helped in planning the size of major subsystem components, as well as the trajectory between targets.

Initialization of Formation Flying Using Primer Vector Theory

NASA Technical Reports Server (NTRS)

Mailhe, Laurie; Schiff, Conrad; Folta, David

2002-01-01

In this paper, we extend primer vector analysis to formation flying. Optimization of the classical rendezvous or free-time transfer problem between two orbits using primer vector theory has been extensively studied for one spacecraft. However, an increasing number of missions are now considering flying a set of spacecraft in close formation. Missions such as the Magnetospheric MultiScale (MMS) and Leonardo-BRDF (Bidirectional Reflectance Distribution Function) need to determine strategies to transfer each spacecraft from the common launch orbit to their respective operational orbit. In addition, all the spacecraft must synchronize their states so that they achieve the same desired formation geometry over each orbit. This periodicity requirement imposes constraints on the boundary conditions that can be used for the primer vector algorithm. In this work we explore the impact of the periodicity requirement in optimizing each spacecraft transfer trajectory using primer vector theory. We first present our adaptation of primer vector theory to formation flying. Using this method, we then compute the AV budget for each spacecraft subject to different formation endpoint constraints.

Inverse simulation system for evaluating handling qualities during rendezvous and docking

NASA Astrophysics Data System (ADS)

Zhou, Wanmeng; Wang, Hua; Thomson, Douglas; Tang, Guojin; Zhang, Fan

2017-08-01

The traditional method used for handling qualities assessment of manned space vehicles is too time-consuming to meet the requirements of an increasingly fast design process. In this study, a rendezvous and docking inverse simulation system to assess the handling qualities of spacecraft is proposed using a previously developed model-predictive-control architecture. By considering the fixed discrete force of the thrusters of the system, the inverse model is constructed using the least squares estimation method with a hyper-ellipsoidal restriction, the continuous control outputs of which are subsequently dispersed by pulse width modulation with sensitivity factors introduced. The inputs in every step are deemed constant parameters, and the method could be considered as a general method for solving nominal, redundant, and insufficient inverse problems. The rendezvous and docking inverse simulation is applied to a nine-degrees-of-freedom platform, and a novel handling qualities evaluation scheme is established according to the operation precision and astronauts' workload. Finally, different nominal trajectories are scored by the inverse simulation and an established evaluation scheme. The scores can offer theoretical guidance for astronaut training and more complex operation missions.

JPRS report: Science and technology. Central Eurasia: Space

NASA Astrophysics Data System (ADS)

1994-12-01

Translated articles cover the following topics: plasma instabilities and space vehicles, need discussed for protection against space catastrophes, Russians offer new energy concept for space stations, Russian space projects: Martian research, multi-impulse rendezvous trajectory for two spacecraft in circular orbit, placement of spacecraft into orbit around Mars with aerobraking, model of the shielding for the inhabited compartments of the base module of the Mir station, and measurement of the background electrostatic and variable electric fields on the outer surface of the Kvant module of the Mir orbital station. There are 25 translated articles in this 28 December 1994 edition.

KSC-2012-1866

NASA Image and Video Library

2012-02-17

Apollo-Soyuz Test Project: The first international crewed spaceflight was a joint U.S.-U.S.S.R. rendezvous and docking mission. The Apollo-Soyuz Test Project, or ASTP, took its name from the spacecraft employed: the American Apollo and the Soviet Soyuz. The three-man Apollo crew lifted off from Kennedy Space Center aboard a Saturn IB rocket on July 15, 1975, to link up with the Soyuz that had launched a few hours earlier. A cylindrical docking module served as an airlock between the two spacecraft for transfer of the crew members. Poster designed by Kennedy Space Center Graphics Department/Greg Lee. Credit: NASA

Antares Post Launch Press Conference

NASA Image and Video Library

2013-09-18

Josh Byerly, public affairs officer, NASA, left, Robert Lightfoot, associate administrator, NASA, second from left, Alan Lindenmoyer, program manager, NASA's Commercial Crew and Cargo Program, and, Frank Culbertson, executive vice president, Orbital Sciences Corporation, right, are seen during a press conference held after the successful launch of the Antares rocket, with the Cygnus cargo spacecraft aboard, Wednesday, Sept. 18, 2013, NASA Wallops Flight Facility, Virginia. Cygnus is on its way to rendezvous with the space station. The spacecraft will deliver about 1,300 pounds (589 kilograms) of cargo, including food and clothing, to the Expedition 37 crew. Photo Credit: (NASA/Bill Ingalls)

Experimental Demonstration of Technologies for Autonomous On-Orbit Robotic Assembly

NASA Technical Reports Server (NTRS)

LeMaster, Edward A.; Schaechter, David B.; Carrington, Connie K.

2006-01-01

The Modular Reconfigurable High Energy (MRHE) program aimed to develop technologies for the automated assembly and deployment of large-scale space structures and aggregate spacecraft. Part of the project involved creation of a terrestrial robotic testbed for validation and demonstration of these technologies and for the support of future development activities. This testbed was completed in 2005, and was thereafter used to demonstrate automated rendezvous, docking, and self-assembly tasks between a group of three modular robotic spacecraft emulators. This paper discusses the rationale for the MRHE project, describes the testbed capabilities, and presents the MRHE assembly demonstration sequence.

Fuel optimal maneuvers of spacecraft about a circular orbit

NASA Technical Reports Server (NTRS)

Carter, T. E.

1982-01-01

Fuel optimal maneuvers of spacecraft relative to a body in circular orbit are investigated using a point mass model in which the magnitude of the thrust vector is bounded. All nonsingular optimal maneuvers consist of intervals of full thrust and coast and are found to contain at most seven such intervals in one period. Only four boundary conditions where singular solutions occur are possible. Computer simulation of optimal flight path shapes and switching functions are found for various boundary conditions. Emphasis is placed on the problem of soft rendezvous with a body in circular orbit.

Pursuit/evasion in orbit

NASA Technical Reports Server (NTRS)

Kelley, H. J.; Cliff, E. M.; Lutze, F. H.

1981-01-01

Maneuvers available to a spacecraft having sufficient propellant to escape an antisatellite satellite (ASAT) attack are examined. The ASAT and the evading spacecraft are regarded as being in circular orbits, and equations of motion are developed for the ASAT to commence a two-impulse maneuver sequence. The ASAT employs thrust impulses which yield a minimum-time-to-rendezvous, considering available fuel. Optimal evasion is shown to involve only in-plane maneuvers, and begins as soon as the ASAT launch information is gathered and thrust activation can be initiated. A closest approach, along with a maximum evasion by the target spacecraft, is calculated to be 14,400 ft. Further research to account for ASATs in parking orbit and for generalization of a continuous control-modeled differential game is indicated.

Multi-rendezvous low-thrust trajectory optimization using costate transforming and homotopic approach

NASA Astrophysics Data System (ADS)

Chen, Shiyu; Li, Haiyang; Baoyin, Hexi

2018-06-01

This paper investigates a method for optimizing multi-rendezvous low-thrust trajectories using indirect methods. An efficient technique, labeled costate transforming, is proposed to optimize multiple trajectory legs simultaneously rather than optimizing each trajectory leg individually. Complex inner-point constraints and a large number of free variables are one main challenge in optimizing multi-leg transfers via shooting algorithms. Such a difficulty is reduced by first optimizing each trajectory leg individually. The results may be, next, utilized as an initial guess in the simultaneous optimization of multiple trajectory legs. In this paper, the limitations of similar techniques in previous research is surpassed and a homotopic approach is employed to improve the convergence efficiency of the shooting process in multi-rendezvous low-thrust trajectory optimization. Numerical examples demonstrate that newly introduced techniques are valid and efficient.

Modular, Reconfigurable, High-Energy Technology Development

NASA Technical Reports Server (NTRS)

Carrington, Connie; Howell, Joe

2006-01-01

The Modular, Reconfigurable High-Energy (MRHE) Technology Demonstrator project was to have been a series of ground-based demonstrations to mature critical technologies needed for in-space assembly of a highpower high-voltage modular spacecraft in low Earth orbit, enabling the development of future modular solar-powered exploration cargo-transport vehicles and infrastructure. MRHE was a project in the High Energy Space Systems (HESS) Program, within NASA's Exploration Systems Research and Technology (ESR&T) Program. NASA participants included Marshall Space Flight Center (MSFC), the Jet Propulsion Laboratory (JPL), and Glenn Research Center (GRC). Contractor participants were the Boeing Phantom Works in Huntsville, AL, Lockheed Martin Advanced Technology Center in Palo Alto, CA, ENTECH, Inc. in Keller, TX, and the University of AL Huntsville (UAH). MRHE's technical objectives were to mature: (a) lightweight, efficient, high-voltage, radiation-resistant solar power generation (SPG) technologies; (b) innovative, lightweight, efficient thermal management systems; (c) efficient, 100kW-class, high-voltage power delivery systems from an SPG to an electric thruster system; (d) autonomous rendezvous and docking technology for in-space assembly of modular, reconfigurable spacecraft; (e) robotic assembly of modular space systems; and (f) modular, reconfigurable distributed avionics technologies. Maturation of these technologies was to be implemented through a series of increasingly-inclusive laboratory demonstrations that would have integrated and demonstrated two systems-of-systems: (a) the autonomous rendezvous and docking of modular spacecraft with deployable structures, robotic assembly, reconfiguration both during assembly and (b) the development and integration of an advanced thermal heat pipe and a high-voltage power delivery system with a representative lightweight high-voltage SPG array. In addition, an integrated simulation testbed would have been developed containing software models representing the technologies being matured in the laboratory demos. The testbed would have also included models for non-MRHE developed subsystems such as electric propulsion, so that end-to-end performance could have been assessed. This paper presents an overview of the MRHE Phase I activities at MSFC and its contractor partners. One of the major Phase I accomplishments is the assembly demonstration in the Lockheed Martin Advanced Technology Center (LMATC) Robot-Satellite facility, in which three robot-satellites successfully demonstrated rendezvous & docking, self-assembly, reconfiguration, adaptable GN&C, deployment, and interfaces between modules. Phase I technology maturation results from ENTECH include material recommendations for radiation hardened Stretched Lens Array (SLA) concentrator lenses, and a design concept and test results for a hi-voltage PV receiver. UAH's accomplishments include Supertube heatpipe test results, which support estimates of thermal conductivities at 30,000 times that of an equivalent silver rod. MSFC performed systems trades and developed a preliminary concept design for a 100kW-class modular reconfigurable solar electric propulsion transport vehicle, and Boeing Phantom Works in Huntsville performed assembly and rendezvous and docking trades. A concept animation video was produced by SAIC, wllich showed rendezvous and docking and SLA-square-rigger deployment in LEO.

Launch Vehicles

NASA Image and Video Library

2004-04-15

The Titan II liftoff. The Titan II launch vehicle was used for carrying astronauts on the Gemini mission. The Gemini Program was an intermediate step between the Project Mercury and the Apollo Program. The major objectives were to subject are two men and supporting equipment to long duration flights, to effect rendezvous and docking with other orbiting vehicle, and to perfect methods of reentry, and landing the spacecraft.

Apollo 7/S-IVB Rendezvous in space

NASA Technical Reports Server (NTRS)

1968-01-01

The expended Saturn IVB stage as photographed from the Apollo 7 spacecraft during transposition and docking maneuvers. This photograph was taken over Sonora, Mexico, during Apollo 7's second revolution of the Earth. The round, white disc inside the open panels of the Saturn IVB is a simulated docking target similar to that used on the lunar module for docking during lunar missions.

ART CONCEPTS - APOLLO IX

NASA Image and Video Library

1969-02-20

S69-19796 (February 1969) --- Composite of six artist's concepts illustrating key events, tasks and activities on the fifth day of the Apollo 9 mission, including vehicles undocked, Lunar Module burns for rendezvous, maximum separation, ascent propulsion system burn, formation flying and docking, and Lunar Module jettison ascent burn. The Apollo 9 mission will evaluate spacecraft lunar module systems performance during manned Earth-orbital flight.

Flyby Characterization of Lower-Degree Spherical Harmonics Around Small Bodies

NASA Technical Reports Server (NTRS)

Takahashi, Yu; Broschart, Stephen; Lantoine, Gregory

2014-01-01

Interest in studying small bodies has grown significantly in the last two decades, and there are a number of past, present, and future missions. These small body missions challenge navigators with significantly different kinds of problems than the planets and moons do. The small bodies' shape is often irregular and their gravitational field significantly weak, which make the designing of a stable orbit a complex dynamical problem. In the initial phase of spacecraft rendezvous with a small body, the determination of the gravitational parameter and lower-degree spherical harmonics are of crucial importance for safe navigation purposes. This motivates studying how well one can determine the total mass and lower-degree spherical harmonics in a relatively short time in the initial phase of the spacecraft rendezvous via flybys. A quick turnaround for the gravity data is of high value since it will facilitate the subsequent mission design of the main scientific observation campaign. We will present how one can approach the problem to determine a desirable flyby geometry for a general small body. We will work in the non-dimensional formulation since it will generalize our results across different size/mass bodies and the rotation rate for a specific combination of gravitational coefficients.

Multi-Sensor Testing for Automated Rendezvous and Docking Sensor Testing at the Flight Robotics Lab

NASA Technical Reports Server (NTRS)

Brewster, Linda L.; Howard, Richard T.; Johnston, A. S.; Carrington, Connie; Mitchell, Jennifer D.; Cryan, Scott P.

2008-01-01

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as AR&D). The crewed missions may also perform rendezvous and docking operations and may require different levels of automation and/or autonomy, and must provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success ofthe Exploration Program. NASA has the responsibility to determine whether the Crew Exploration Vehicle (CEV) contractor-proposed relative navigation sensor suite will meet the requirements. The relatively low technology readiness level of AR&D relative navigation sensors has been carried as one of the CEV Project's top risks. The AR&D Sensor Technology Project seeks to reduce the risk by the testing and analysis of selected relative navigation sensor technologies through hardware-in-the-Ioop testing and simulation. These activities will provide the CEV Project information to assess the relative navigation sensors maturity as well as demonstrate test methods and capabilities. The first year of this project focused on a series of "pathfinder" testing tasks to develop the test plans, test facility requirements, trajectories, math model architecture, simulation platform, and processes that will be used to evaluate the Contractor-proposed sensors. Four candidate sensors were used in the first phase of the testing. The second phase of testing used four sensors simultaneously: two Marshall Space Flight Center (MSFC) Advanced Video Guidance Sensors (AVGS), a laser-based video sensor that uses retroreflectors attached to the target vehicle, and two commercial laser range finders. The multi-sensor testing was conducted at MSFC's Flight Robotics Laboratory (FRL) using the FRL's 6-DOF gantry system, called the Dynamic Overhead Target System (DOTS). The target vehicle for "docking" in the laboratory was a mockup that was representative of the proposed CEV docking system, with added retroreflectors for the AVGS.' The multi-sensor test configuration used 35 open-loop test trajectories covering three major objectives: (l) sensor characterization trajectories designed to test a wide range of performance parameters; (2) CEV-specific trajectories designed to test performance during CEV-like approach and departure profiles; and (3) sensor characterization tests designed for evaluating sensor performance under more extreme conditions as might be induced during a spacecraft failure or during contingency situations. This paper describes the test development, test facility, test preparations, test execution, and test results of the multisensor series oftrajectories

A hybrid systems strategy for automated spacecraft tour design and optimization

NASA Astrophysics Data System (ADS)

Stuart, Jeffrey R.

As the number of operational spacecraft increases, autonomous operations is rapidly evolving into a critical necessity. Additionally, the capability to rapidly generate baseline trajectories greatly expands the range of options available to analysts as they explore the design space to meet mission demands. Thus, a general strategy is developed, one that is suitable for the construction of flight plans for both Earth-based and interplanetary spacecraft that encounter multiple objects, where these multiple encounters comprise a ``tour''. The proposed scheme is flexible in implementation and can readily be adjusted to a variety of mission architectures. Heuristic algorithms that autonomously generate baseline tour trajectories and, when appropriate, adjust reference solutions in the presence of rapidly changing environments are investigated. Furthermore, relative priorities for ranking the targets are explicitly accommodated during the construction of potential tour sequences. As a consequence, a priori, as well as newly acquired, knowledge concerning the target objects enhances the potential value of the ultimate encounter sequences. A variety of transfer options are incorporated, from rendezvous arcs enabled by low-thrust engines to more conventional impulsive orbit adjustments via chemical propulsion technologies. When advantageous, trajectories are optimized in terms of propellant consumption via a combination of indirect and direct methods; such a combination of available technologies is an example of hybrid optimization. Additionally, elements of hybrid systems theory, i.e., the blending of dynamical states, some discrete and some continuous, are integrated into the high-level tour generation scheme. For a preliminary investigation, this strategy is applied to mission design scenarios for a Sun-Jupiter Trojan asteroid tour as well as orbital debris removal for near-Earth applications.

Artist's Concept of the Apollo-Soyuz Test Project

NASA Technical Reports Server (NTRS)

1974-01-01

This artist's concept depicts the Apollo-Soyuz Test Project (ASTP), the first international docking of the U.S.'s Apollo spacecraft and the U.S.S.R.'s Soyuz spacecraft in space. The objective of the ASTP mission was to provide the basis for a standardized international system for docking of marned spacecraft. The Soyuz spacecraft, with Cosmonauts Alexei Leonov and Valeri Kubasov aboard, was launched from the Baikonur Cosmodrome near Tyuratam in the Kazakh, Soviet Socialist Republic, at 8:20 a.m. (EDT) on July 15, 1975. The Apollo spacecraft, with Astronauts Thomas Stafford, Vance Brand, and Donald Slayton aboard, was launched from Launch Complex 39B, Kennedy Space Center, Florida, at 3:50 p.m. (EDT) on July 15, 1975. The Primary objectives of the ASTP were achieved. They performed spacecraft rendezvous, docking and undocking, conducted intervehicular crew transfer, and demonstrated the interaction of U.S. and U.S.S.R. control centers and spacecraft crews. The mission marked the last use of a Saturn launch vehicle. The Marshall Space Flight Center was responsible for development and sustaining engineering of the Saturn IB launch vehicle during the mission.

Asteroid Redirect Crewed Mission Nominal Design and Performance

NASA Technical Reports Server (NTRS)

Condon, Gerald; williams, Jacob

2014-01-01

In 2010, the President announced that, in 2025, the U.S. intended to launch a human mission to an asteroid [1]. This announcement was followed by the idea of a Capability Driven Framework (CDF) [2], which is based on the idea of evolving capabilities from less demanding to more demanding missions to multiple possible destinations and with increased flexibility, cost effectiveness and sustainability. Focused missions, such as a NASA inter-Center study that examined the viability and implications of sending a crew to a Near Earth Asteroid (NEA) [3], provided a way to better understand and evaluate the utility of these CDF capabilities when applied to an actual mission. The long duration of the NEA missions were contrasted with a concept described in a study prepared for the Keck Institute of Space Studies (KISS) [4] where a robotic spacecraft would redirect an asteroid to the Earth-Moon vicinity, where a relatively short duration crewed mission could be conducted to the captured asteroid. This mission concept was included in the National Aeronautics and Space Administration (NASA) fiscal year 2014 budget request, as submitted by the NASA Administrator [5]. NASA studies continued to examine the idea of a crewed mission to a captured asteroid in the Earth-Moon vicinity. During this time was an announcement of NASA's Asteroid Grand Challenge [6]. Key goals for the Asteroid Grand Challenge are to locate, redirect, and explore an asteroid, as well as find and plan for asteroid threats. An Asteroid Redirect Mission (ARM) study was being conducted, which supports this Grand Challenge by providing understanding in how to execute an asteroid rendezvous, capture it, and redirect it to Earth-Moon space, and, in particular, to a distant retrograde orbit (DRO). Subsequent to the returning of the asteroid to a DRO, would be the launch of a crewed mission to rendezvous with the redirected asteroid. This report examines that crewed mission by assessing the Asteroid Redirect Crewed Mission (ARCM) nominal design and performance costs associated with an Orion based crewed rendezvous mission to a captured asteroid in an Earth-Moon DRO. The ARM study includes two fundamental mission phases: 1) The Asteroid Redirect Robotic Mission (ARRM) and 2) the ARCM. The ARRM includes a solar electric propulsion based robotic asteroid return vehicle (ARV) sent to rendezvous with a selected near Earth asteroid, capture it, and return it to a DRO in the Earth-Moon vicinity. The DRO is selected over other possible asteroid parking orbits due to its achievability (by both the robotic and crewed vehicles) and by its stability (e.g., no orbit maintenance is required). After the return of the asteroid to the Earth-Moon vicinity, the ARCM is executed and carries a crew of two astronauts to a DRO to rendezvous with the awaiting ARV with the asteroid. The outbound and inbound transfers employ lunar gravity assist (LGA) flybys to reduce the Orion propellant requirement for the overall nominal mission, which provides a nominal mission with some reserve propellant for possible abort situations. The nominal mission described in this report provides a better understanding of the mission considerations as well as the feasibility of such a crewed mission, particularly with regard to spacecraft currently undergoing development, such as the Orion vehicle and the Space Launch System (SLS).

Characterization of the Surface Properties of MUSES-C/Hayabusa Spacecraft Target Asteroid 25143 Itokawa (1998 SF36)

NASA Technical Reports Server (NTRS)

Lederer, S. M.; Domingue, D. L.; Vilas, F.; Abe, M.; Farnham, T. L.; Jarvis, K. S.; Lowry, S. C.; Ohba, Y.; Weissman, P. R.; French, L. M.

2004-01-01

Several spacecraft missions have recently targeted asteroids to study their morphologies and physical properties (e.g. Galileo, NEAR Shoemaker), and more are planned. MUSES-C is a Japanese mission designed to rendezvous with a near-Earth asteroid (NEA). The MUSES-C spacecraft, Hayabusa, was launched successfully in May 2003. It will rendezvous with its target asteroid in 2005, and return samples to the Earth in 2007. Its target, 25143 Itokawa (1998 SF36), made a close approach to the Earth in 2001. We collected an extensive ground-based database of broadband photometry obtained during this time, which maximized the phase angle coverage, to characterize this target in preparation for the mission. Our project was designed to capitalize on the broadband UBVRI photometric observations taken with a series of telescopes, instrumentation, and observers. Photometry and spectrophotometry of Itokawa were acquired at Lowell, McDonald, Steward, Palomar, Table Mountain and Kiso Observatories. The photometric data sets were combined to calculate Hapke model parameters of the surface material of Itokawa, and examine the solar-corrected broadband color characteristics of the asteroid. Broadband photometry of an object can be used to: (1) determine its colors and thereby contribute to the understanding of its surface composition and taxonomic class, and (2) infer global physical surface properties of the target body. We present both colors from UBVRI observations of the MUSES-C target Itokawa, and physical properties derived by applying a Hapke model to the broadband BVRI photometry.

A new approach to impulsive rendezvous near circular orbit

NASA Astrophysics Data System (ADS)

Carter, Thomas; Humi, Mayer

2012-04-01

A new approach is presented for the problem of planar optimal impulsive rendezvous of a spacecraft in an inertial frame near a circular orbit in a Newtonian gravitational field. The total characteristic velocity to be minimized is replaced by a related characteristic-value function and this related optimization problem can be solved in closed form. The solution of this problem is shown to approach the solution of the original problem in the limit as the boundary conditions approach those of a circular orbit. Using a form of primer-vector theory the problem is formulated in a way that leads to relatively easy calculation of the optimal velocity increments. A certain vector that can easily be calculated from the boundary conditions determines the number of impulses required for solution of the optimization problem and also is useful in the computation of these velocity increments. Necessary and sufficient conditions for boundary conditions to require exactly three nonsingular non-degenerate impulses for solution of the related optimal rendezvous problem, and a means of calculating these velocity increments are presented. A simple example of a three-impulse rendezvous problem is solved and the resulting trajectory is depicted. Optimal non-degenerate nonsingular two-impulse rendezvous for the related problem is found to consist of four categories of solutions depending on the four ways the primer vector locus intersects the unit circle. Necessary and sufficient conditions for each category of solutions are presented. The region of the boundary values that admit each category of solutions of the related problem are found, and in each case a closed-form solution of the optimal velocity increments is presented. Similar results are presented for the simpler optimal rendezvous that require only one-impulse. For brevity degenerate and singular solutions are not discussed in detail, but should be presented in a following study. Although this approach is thought to provide simpler computations than existing methods, its main contribution may be in establishing a new approach to the more general problem.

Spheres: from Ground Development to ISS Operations

NASA Technical Reports Server (NTRS)

Katterhagen, A.

2016-01-01

SPHERES (Synchronized Position Hold Engage and Reorient Experimental Satellites) is an internal International Space Station (ISS) Facility that supports multiple investigations for the development of multi-spacecraft and robotic control algorithms. The SPHERES National Lab Facility aboard ISS is managed and operated by NASA Ames Research Center (ARC) at Moffett Field California. The SPHERES Facility on ISS consists of three self-contained eight-inch diameter free-floating satellites which perform the various flight algorithms and serve as a platform to support the integration of experimental hardware. SPHERES has served to mature the adaptability of control algorithms of future formation flight missions in microgravity (6 DOF (Degrees of Freedom) / long duration microgravity), demonstrate key close-proximity formation flight and rendezvous and docking maneuvers, understand fault diagnosis and recovery, improve the field of human telerobotic operation and control, and lessons learned on ISS have significant impact on ground robotics, mapping, localization, and sensing in three-dimensions - among several other areas of study.

Low-thrust trajectory optimization of asteroid sample return mission with multiple revolutions and moon gravity assists

NASA Astrophysics Data System (ADS)

Tang, Gao; Jiang, FanHuag; Li, JunFeng

2015-11-01

Near-Earth asteroids have gained a lot of interest and the development in low-thrust propulsion technology makes complex deep space exploration missions possible. A mission from low-Earth orbit using low-thrust electric propulsion system to rendezvous with near-Earth asteroid and bring sample back is investigated. By dividing the mission into five segments, the complex mission is solved separately. Then different methods are used to find optimal trajectories for every segment. Multiple revolutions around the Earth and multiple Moon gravity assists are used to decrease the fuel consumption to escape from the Earth. To avoid possible numerical difficulty of indirect methods, a direct method to parameterize the switching moment and direction of thrust vector is proposed. To maximize the mass of sample, optimal control theory and homotopic approach are applied to find the optimal trajectory. Direct methods of finding proper time to brake the spacecraft using Moon gravity assist are also proposed. Practical techniques including both direct and indirect methods are investigated to optimize trajectories for different segments and they can be easily extended to other missions and more precise dynamic model.

The Near Earth Object (NEO) Scout Spacecraft: A Low-cost Approach to In-situ Characterization of the NEO Population

NASA Technical Reports Server (NTRS)

Woeppel, Eric A.; Balsamo, James M.; Fischer, Karl J.; East, Matthew J.; Styborski, Jeremy A.; Roche, Christopher A.; Ott, Mackenzie D.; Scorza, Matthew J.; Doherty, Christopher D.; Trovato, Andrew J.;

Adaptive relative pose control for autonomous spacecraft rendezvous and proximity operations with thrust misalignment and model uncertainties

NASA Astrophysics Data System (ADS)

Sun, Liang; Zheng, Zewei

2017-04-01

An adaptive relative pose control strategy is proposed for a pursue spacecraft in proximity operations on a tumbling target. Relative position vector between two spacecraft is required to direct towards the docking port of the target while the attitude of them must be synchronized. With considering the thrust misalignment of pursuer, an integrated controller for relative translational and relative rotational dynamics is developed by using norm-wise adaptive estimations. Parametric uncertainties, unknown coupled dynamics, and bounded external disturbances are compensated online by adaptive update laws. It is proved via Lyapunov stability theory that the tracking errors of relative pose converge to zero asymptotically. Numerical simulations including six degrees-of-freedom rigid body dynamics are performed to demonstrate the effectiveness of the proposed controller.

LIFTOFF - GEMINI-TITAN (GT)-9A - ATLAS/AGENA - CAPE

NASA Image and Video Library

1966-05-17

S66-34610 (17 May 1966) --- An Agena Target Vehicle atop its Atlas Launch vehicle is launched from the Kennedy Space Center (KSC) Launch Complex 14 at 10:15 am., May 17, 1966. The Agena was intended as a rendezvous and docking vehicle for the Gemini-9 spacecraft. However, since the Agena failed to achieve orbit, the Gemini-9 mission was postponed. Photo credit: NASA

REVIEW GT-7 MISSION REQUIREMENTS (PRIME CREW)

NASA Image and Video Library

1965-12-02

S65-56313 (2 Dec. 1965) --- Gemini-7 astronauts James Lovell Jr. (center) and Frank Borman (right) review mission requirements for their Gemini-7 flight. The two astronauts are scheduled for a 14-day mission. On the fifth day, they will attempt a rendezvous with the Gemini-6 spacecraft to be launched nine days later and piloted by astronaut Walter Schirra Jr. and Thomas Stafford. Photo credit: NASA

Progress 47P Redocking

NASA Image and Video Library

2012-07-29

ISS032-E-010629 (28 July 2012) --- The unpiloted Russian Progress 47 resupply spacecraft is featured in this image photographed by an Expedition 32 crew member as it approaches the International Space Station for re-docking on July 28, 2012. The cargo ship temporarily undocked on July 22 in order to test the new Kurs-NA automated rendezvous system. Station solar array panels are visible in the foreground.

Evaluating small-body landing hazards due to blocks

NASA Astrophysics Data System (ADS)

Ernst, C.; Rodgers, D.; Barnouin, O.; Murchie, S.; Chabot, N.

2014-07-01

Introduction: Landed missions represent a vital stage of spacecraft exploration of planetary bodies. Landed science allows for a wide variety of measurements essential to unraveling the origin and evolution of a body that are not possible remotely, including but not limited to compositional measurements, microscopic grain characterization, and the physical properties of the regolith. To date, two spacecraft have performed soft landings on the surface of a small body. In 2001, the Near Earth Asteroid Rendezvous (NEAR) mission performed a controlled descent and landing on (433) Eros following the completion of its mission [1]; in 2005, the Hayabusa spacecraft performed two touch-and-go maneuvers at (25143) Itokawa [2]. Both landings were preceded by rendezvous spacecraft reconnaissance, which enabled selection of a safe landing site. Three current missions have plans to land on small bodies (Rosetta, Hayabusa 2, and OSIRIS-REx); several other mission concepts also include small-body landings. Small-body landers need to land at sites having slopes and block abundances within spacecraft design limits. Due to the small scale of the potential hazards, it can be difficult or impossible to fully characterize a landing surface before the arrival of the spacecraft at the body. Although a rendezvous mission phase can provide global reconnaissance from which a landing site can be chosen, reasonable a priori assurance that a safe landing site exists is needed to validate the design approach for the spacecraft. Method: Many robotic spacecraft have landed safely on the Moon and Mars. Images of these landing sites, as well as more recent, extremely high-resolution orbital datasets, have enabled the comparison of orbital block observations to the smaller blocks that pose hazards to landers. Analyses of the Surveyor [3], Viking 1 and 2, Mars Pathfinder, Phoenix, Spirit, Opportunity, and Curiosity landing sites [4--8] have indicated that for a reasonable difference in size (a factor of several to ten), the size-frequency distribution of blocks can be modeled, allowing extrapolation from large block distributions to estimate small block densities. From that estimate, the probability of a lander encountering hazardous blocks can be calculated for a given lander design. Such calculations are used routinely to vet candidate sites for Mars landers [5--8]. Application to Small Bodies: To determine whether a similar approach will work for small bodies, we must determine if the large and small block populations can be linked. To do so, we analyze the comprehensive block datasets for the intermediate-sized Eros [9,10] and the small Itokawa [11,12]. Global and local block size-frequency distributions for Eros and Itokawa have power-law slopes on the order of -3 and match reasonably well between larger block sizes (from lower-resolution images) and smaller block sizes (from higher-resolution images). Although absolute block densities differ regionally on each asteroid, the slopes match reasonably well between Itokawa and Eros, with the geologic implications of this result discussed in [10]. For Eros and Itokawa, the approach of extending the size-frequency distribution from large, tens-of-meter-sized blocks down to small, tens-of-centimeter-sized blocks using a power-law fit to the large population yields reasonable estimates of small block populations. It is important to note that geologic context matters for the absolute block density --- if the global counts include multiple geologic settings, they will not directly extend to local areas containing only one setting [10]. A small number of high-resolution images of Phobos are sufficient for measuring blocks. These images are concentrated in the area outside of Stickney crater, which is thought to be the source of most of the observed blocks [13]. Block counts by Thomas et al. [13] suggest a power-law slope similar to those of Eros [9] and Itokawa global counts, with the absolute density of blocks similar to that of global Eros. Because blocks tend to be more numerous proximal to large, young craters (e.g., Stickney on Phobos, Shoemaker on Eros), the block density across most of Phobos is likely to be lower than that observed in the available high-resolution images. We suggest that a power-law extrapolation of Eros or Phobos large-block distributions provides upper limits for assessing the block landing hazards faced by a Phobos lander.

Advanced Navigation Strategies For Asteroid Sample Return Missions

NASA Technical Reports Server (NTRS)

Getzandanner, K.; Bauman, J.; Williams, B.; Carpenter, J.

2010-01-01

Flyby and rendezvous missions to asteroids have been accomplished using navigation techniques derived from experience gained in planetary exploration. This paper presents analysis of advanced navigation techniques required to meet unique challenges for precision navigation to acquire a sample from an asteroid and return it to Earth. These techniques rely on tracking data types such as spacecraft-based laser ranging and optical landmark tracking in addition to the traditional Earth-based Deep Space Network radio metric tracking. A systematic study of navigation strategy, including the navigation event timeline and reduction in spacecraft-asteroid relative errors, has been performed using simulation and covariance analysis on a representative mission.

KSC-99pc40

NASA Image and Video Library

1999-01-11

Workers in the Payload Hazardous Servicing Facility watch as the Stardust spacecraft is rotated and lowered before deploying the solar panels for lighting tests. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule (seen on top of the spacecraft) to be jettisoned as it swings by Earth in January 2006

Potential applications of MMC and aluminum-lithium alloys in cameras for CRAF spacecraft. [Comet Rendezvous Asteroid Flyby Mission

NASA Technical Reports Server (NTRS)

Lane, Marc; Hsieh, Cheng; Adams, Lloyd

1989-01-01

In undertaking the design of a 2000-mm focal length camera for the Mariner Mark II series of spacecraft, JPL sought novel materials with the requisite dimensional and thermal stability, outgassing and corrosion resistance, low mass, high stiffness, and moderate cost. Metal-matrix composites and Al-Li alloys have, in addition to excellent mechanical properties and low density, a suitably low coefficient of thermal expansion, high specific stiffness, and good electrical conductivity. The greatest single obstacle to application of these materials to camera structure design is noted to have been the lack of information regarding long-term dimensional stability.

Design of Spacecraft Missions to Remove Multiple Orbital Debris Objects

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Alfano, Salvatore; Pinon, Elfego; Gold, Kenn; Gaylor, David

2012-01-01

The amount of hazardous debris in Earth orbit has been increasing, posing an evergreater danger to space assets and human missions. In January of 2007, a Chinese ASAT test produced approximately 2600 pieces of orbital debris. In February of 2009, Iridium 33 collided with an inactive Russian satellite, yielding approximately 1300 pieces of debris. These recent disastrous events and the sheer size of the Earth orbiting population make clear the necessity of removing orbital debris. In fact, experts from both NASA and ESA have stated that 10 to 20 pieces of orbital debris need to be removed per year to stabilize the orbital debris environment. However, no spacecraft trajectories have yet been designed for removing multiple debris objects and the size of the debris population makes the design of such trajectories a daunting task. Designing an efficient spacecraft trajectory to rendezvous with each of a large number of orbital debris pieces is akin to the famous Traveling Salesman problem, an NP-complete combinatorial optimization problem in which a number of cities are to be visited in turn. The goal is to choose the order in which the cities are visited so as to minimize the total path distance traveled. In the case of orbital debris, the pieces of debris to be visited must be selected and ordered such that spacecraft propellant consumption is minimized or at least kept low enough to be feasible. Emergent Space Technologies, Inc. has developed specialized algorithms for designing efficient tour missions for near-Earth asteroids that may be applied to the design of efficient spacecraft missions capable of visiting large numbers of orbital debris pieces. The first step is to identify a list of high priority debris targets using the Analytical Graphics, Inc. SOCRATES website and then obtain their state information from Celestrak. The tour trajectory design algorithms will then be used to determine the itinerary of objects and v requirements. These results will shed light on how many debris pieces can be visited for various amounts of propellant, which launch vehicles can accommodate such missions, and how much margin is available for debris removal system payloads.

Progress 33P undock

NASA Image and Video Library

2009-06-30

ISS020-E-015987 (30 June 2009) --- An unpiloted ISS Progress 33 cargo craft, filled with trash and unneeded items, departs from the International Space Station?s Pirs Docking Compartment at 1:30 p.m. (CDT) on June 30, 2009. The Progress was commanded into a parking orbit for its re-rendezvous with the ISS on July 12, approaching to within 10-15 meters of the Zvezda Service Module to test new automated rendezvous equipment mounted on Zvezda during a pair of spacewalks earlier this month by Gennady Padalka and Mike Barratt that will be used to guide the new Mini-Research Module-2 (MRM2) to an unpiloted docking to the zenith port of Zvezda later this year. MRM2 will serve as a new docking port for Russian spacecraft and an additional airlock for spacewalks conducted out of the Russian segment.

Pose estimation and tracking of non-cooperative rocket bodies using Time-of-Flight cameras

NASA Astrophysics Data System (ADS)

Gómez Martínez, Harvey; Giorgi, Gabriele; Eissfeller, Bernd

2017-10-01

This paper presents a methodology for estimating the position and orientation of a rocket body in orbit - the target - undergoing a roto-translational motion, with respect to a chaser spacecraft, whose task is to match the target dynamics for a safe rendezvous. During the rendezvous maneuver the chaser employs a Time-of-Flight camera that acquires a point cloud of 3D coordinates mapping the sensed target surface. Once the system identifies the target, it initializes the chaser-to-target relative position and orientation. After initialization, a tracking procedure enables the system to sense the evolution of the target's pose between frames. The proposed algorithm is evaluated using simulated point clouds, generated with a CAD model of the Cosmos-3M upper stage and the PMD CamCube 3.0 camera specifications.

Autonomous Deep-Space Optical Navigation Project

NASA Technical Reports Server (NTRS)

D'Souza, Christopher

2014-01-01

This project will advance the Autonomous Deep-space navigation capability applied to Autonomous Rendezvous and Docking (AR&D) Guidance, Navigation and Control (GNC) system by testing it on hardware, particularly in a flight processor, with a goal of limited testing in the Integrated Power, Avionics and Software (IPAS) with the ARCM (Asteroid Retrieval Crewed Mission) DRO (Distant Retrograde Orbit) Autonomous Rendezvous and Docking (AR&D) scenario. The technology, which will be harnessed, is called 'optical flow', also known as 'visual odometry'. It is being matured in the automotive and SLAM (Simultaneous Localization and Mapping) applications but has yet to be applied to spacecraft navigation. In light of the tremendous potential of this technique, we believe that NASA needs to design a optical navigation architecture that will use this technique. It is flexible enough to be applicable to navigating around planetary bodies, such as asteroids.

Apollo-Soyuz Test Project (ASTP)

NASA Technical Reports Server (NTRS)

1974-01-01

This artist's concept depicts the Apollo-Soyuz Test Project (ASTP) with insets of photographs of three U.S. astronauts (Thomas Stafford, Vance Brand, and Donald Slayton) and two U.S.S.R. cosmonauts (Alexei Leonov and Valeri Kubasov). The objective of the ASTP mission was to accomplish the first docking of a standardized international system, the U.S.'s Apollo spacecraft and the U.S.S.R.'s Soyuz spacecraft, in space. The Soyuz spacecraft was launched from the Baikonur Cosmodrome near Tyuratam in the Kazakh, Soviet Socialist Republic, at 8:20 a.m. (EDT) on July 15, 1975. The Apollo spacecraft was launched from Launch Complex 39B, Kennedy Space Center, Florida, at 3:50 p.m. (EDT) on July 15, 1975. The Primary objectives of the ASTP were achieved. They performed spacecraft rendezvous, docking and undocking, conducted intervehicular crew transfer, and demonstrated the interaction of U.S. and U.S.S.R. control centers and spacecraft crews. The mission marked the last use of a Saturn launch vehicle. The Marshall Space Flight Center was responsible for development and sustaining engineering of the Saturn IB launch vehicle during the mission.

Boom Rendezvous Alternative Docking Approach

NASA Technical Reports Server (NTRS)

Bonometti, Joseph A.

2006-01-01

Space rendezvous and docking has always been attempted with primarily one philosophic methodology. The slow matching of one vehicle's orbit by a second vehicle and then a final closing sequence that ends in matching the orbits with perfect precision and with near zero relative velocities. The task is time consuming, propellant intensive, risk inherent (plume impingement, collisions, fuel depletion, etc.) and requires substantial hardware mass. The historical background and rationale as to why this approach is used is discussed in terms of the path-not-taken and in light of an alternate methodology. Rendezvous and docking by boom extension is suggested to have inherent advantages that today s technology can readily exploit. Extension from the primary spacecraft, beyond its inherent large inertia, allows low inertia connections to be made rapidly and safely. Plume contamination issues are eliminated as well as the extra propellant mass and risk required for the final thruster (docking) operations. Space vehicle connection hardware can be significantly lightened. Also, docking sensors and controls require less fidelity; allowing them to be more robust and less sensitive. It is the potential safety advantage and mission risk reduction that makes this approach attractive, besides the prospect of nominal time and mass savings.

COMPASS Final Report: Near Earth Asteroids Rendezvous and Sample Earth Returns (NEARER)

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; McGuire, Melissa L.

2009-01-01

In this study, the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team completed a design for a multi-asteroid (Nereus and 1996 FG3) sample return capable spacecraft for the NASA In-Space Propulsion Office. The objective of the study was to support technology development and assess the relative benefits of different electric propulsion systems on asteroid sample return design. The design uses a single, heritage Orion solar array (SA) (approx.6.5 kW at 1 AU) to power a single NASA Evolutionary Xenon Thruster ((NEXT) a spare NEXT is carried) to propel a lander to two near Earth asteroids. After landing and gathering science samples, the Solar Electric Propulsion (SEP) vehicle spirals back to Earth where it drops off the first sample s return capsule and performs an Earth flyby to assist the craft in rendezvousing with a second asteroid, which is then sampled. The second sample is returned in a similar fashion. The vehicle, dubbed Near Earth Asteroids Rendezvous and Sample Earth Returns (NEARER), easily fits in an Atlas 401 launcher and its cost estimates put the mission in the New Frontier s (NF's) class mission.

Vacuum to Antimatter-Rocket Interstellar Explorer System (VARIES): A Proposed Program for an Interstellar Rendezvous and Return Architecture

NASA Astrophysics Data System (ADS)

Obousy, R.

While interstellar missions have been explored in the literature, one mission architecture has not received much attention, namely the interstellar rendezvous and return mission that could be accomplished on timescales comparable with a working scientist's career. Such a mission would involve an initial boost phase followed by a coasting phase to the target system. Next would be the deceleration and rendezvous phase, which would be followed by a period of scientific data gathering. Finally, there would be a second boost phase, aimed at returning the spacecraft back to the solar system, and subsequent coasting and deceleration phases upon return to our solar system. Such a mission would represent a precursor to a future manned interstellar mission; which in principle could safely return any astronauts back to Earth. In this paper a novel architecture is proposed that would allow for an unmanned interstellar rendezvous and return mission. The approach utilized for the Vacuum to Antimatter-Rocket Interstellar Explorer System (VARIES) would lead to system components and mission approaches that could be utilized for autonomous operation of other deep-space probes. Engineering solutions for such a mission will have a significant impact on future exploration and sample return missions for the outer planets. This paper introduces the general concept, with a mostly qualitative analysis. However, a full research program is introduced, and as this program progresses, more quantitative papers will be released.

A mission concept for a Grand Tour of Multiple Asteroid Systems

NASA Astrophysics Data System (ADS)

Marchis, F.; Dankanich, J.; Tricarico, P.; Bellerose, J.

2009-12-01

In 1993, the Galileo spacecraft imaged the first companion of asteroid, Dactyl orbiting 243 Ida, a main-belt asteroid. Since then, discoveries have been accumulated thanks to the development of high angular resolution imaging on ground-based telescopes (adaptive optics), radar observations and accurate photometric light curve measurements. To date, 180 companions of small solar system bodies (SSSBs) are known in various populations, including 100 in the asteroid main belt, 33 Near Earth Asteroids, 4 Jupiter-Trojan asteroids and 44 in the Kuiper Belt. Multiple Asteroids have been shown to be complex worlds in their own with a wide range of morphologies, dynamical histories, and structural evolution. To the exception of 243 Ida, no spacecraft has visited any of them. Investigating binary asteroid systems can verify and validate current theories on their formation and on the influence of the sun in their formation (YORP effect) and evolution (space weathering). In particular, assessing the origin of the secondary satellite, if it is of common origin or capture, can provide clue of their formation. To a larger extend, the determination of their nature, scenario formation and evolution are key to understand how planet formation occurred but also to understand i) the population and compositional structure of the SSSB today ii) how the dynamics and collisions modify this structure over time iii) what the physical properties of asteroids are (density, porosity) iv) how the surface modification processes affect our ability to determine this structure (e.g. space weathering). In addition, being able to study these properties on closeby asteroids will give a relative scale accounting for the sizes, shape, rotation periods and cratering rate of these small and young bodies. In the framework of the NASA Discovery program, we propose a mission consisting of a Grand Tour of several multiple asteroid systems, including the flyby of a near earth binary asteroid and the rendezvous with several multiple asteroid systems located in the main belt. This mission concept uses the NASA's evolutionary Xenon Thruster (NEXT), the second generation of electric propulsion with 3 times more input power than the previous generation (NSTAR) of the Dawn mission. The mission objectives for each rendezvous asteroid are i) the characterization of the surface geology by direct imaging in visible and thermal infrared spectroscopy, ii) the characterization of the shape and gravity coupling visible observations with LIDAR ranging data, iii) the determination of the thermophysical properties of the surface, and iv) the identification of the surface composition by visible and near-infrared spectroscopy. The trajectory, science package and mission operations of the mission will be described. This work is supported by the National Science Foundation 05-608, "Astronomy and Astrophysics Research Grants (AAG)" No AST-0807468

Apollo 7/S-IVB Rendezvous in space

NASA Image and Video Library

1968-10-11

AS07-03-1538 (11 Oct. 1968) --- The expended Saturn IVB stage as photographed from the Apollo 7 spacecraft during transposition and docking maneuvers. This photograph was taken during Apollo 7's second revolution of Earth. Earth below has heavy cloud cover. The round, white disc inside the open panels of the Saturn IVB is a simulated docking target similar to that used on the lunar module for docking during lunar missions.

Apollo 7/S-IVB Rendezvous in space

NASA Image and Video Library

1968-10-11

AS07-03-1531 (11 Oct. 1968) --- The expended Saturn IVB stage as photographed from the Apollo 7 spacecraft during transposition and docking maneuvers. This photograph was taken over Sonora, Mexico, during Apollo 7's second revolution of Earth. The round, white disc inside the open panels of the Saturn IVB is a simulated docking target similar to that used on the lunar module for docking during lunar missions.

Optimal Control of a Circular Satellite Formation Subject to Gravitational Perturbations

DTIC Science & Technology

2007-03-01

fundamental reference in the study of the dynamics of close-proximity spacecraft is the paper by Clohessy and Wiltshire (5). In this work, the linear...dynamics for a satellite rendezvous problem are derived, which are now commonly known as either the Clohessy - Wiltshire (CW) equations or Hill’s...themselves to closed-form solutions, as did the Clohessy - Wiltshire development. When the nonlinear approach is undertaken, the numeric integration

A Framework for Designing Optimal Spacecraft Formations

DTIC Science & Technology

2002-09-01

to the Hill- Clohessy - Wiltshire equations were reproduced. For an example using elliptical reference orbits, Reference 17 outlines a solution with...2001. 15. Clohessy , W.H. and Wiltshire , R. S., “Terminal Guidance System for Satellite Rendezvous,” Journal of the Aerospace Sciences, Vol.27, No...Hill- Clohessy -Wiltshire15 (C-W) equations were chosen as the first model specifically because the solutions were known. This allowed a validation

Space Shuttle Familiarization

NASA Technical Reports Server (NTRS)

Mellett, Kevin

2006-01-01

This slide presentation visualizes the NASA space center and research facility sites, as well as the geography, launching sites, launching pads, rocket launching, pre-flight activities, and space shuttle ground operations located at NASA Kennedy Space Center. Additionally, highlights the international involvement behind the International Space Station and the space station mobile servicing system. Extraterrestrial landings, surface habitats and habitation systems, outposts, extravehicular activity, and spacecraft rendezvous with the Earth return vehicle are also covered.

OFFICIAL EMBLEM - APOLLO-SOYUZ TEST PROJECT (ASTP)

NASA Image and Video Library

1974-03-01

S74-17843 (March 1974) --- This is the official emblem of the Apollo-Soyuz Test Project chosen by NASA and the Soviet Academy of Sciences. The joint U.S.-USSR space mission is scheduled to be flown in July 1975. Of circular design, the emblem has the words Apollo in English and Soyuz in Russian around a center disc which depicts the two spacecraft docked together in Earth orbit. The Apollo-Soyuz Test Project will be carried out by a Soviet Soyuz spacecraft and a U.S. Apollo spacecraft which will rendezvous and dock in orbit. Soyuz and Apollo will remain docked for as long as two days in which period, the three Apollo astronauts will enter Soyuz and the two Soyuz cosmonauts will visit Apollo via a docking module. The Russian word "soyuz" means "union" in English.

Multi-Sensor Testing for Automated Rendezvous and Docking Sensor Testing at the Flight Robotics Laboratory

NASA Technical Reports Server (NTRS)

Brewster, L.; Johnston, A.; Howard, R.; Mitchell, J.; Cryan, S.

2007-01-01

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as AR&D). The crewed missions may also perform rendezvous and docking operations and may require different levels of automation and/or autonomy, and must provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success of the Exploration Program. NASA has the responsibility to determine whether the Crew Exploration Vehicle (CEV) contractor proposed relative navigation sensor suite will meet the requirements. The relatively low technology readiness level of AR&D relative navigation sensors has been carried as one of the CEV Project's top risks. The AR&D Sensor Technology Project seeks to reduce the risk by the testing and analysis of selected relative navigation sensor technologies through hardware-in-the-loop testing and simulation. These activities will provide the CEV Project information to assess the relative navigation sensors maturity as well as demonstrate test methods and capabilities. The first year of this project focused on a series of"pathfinder" testing tasks to develop the test plans, test facility requirements, trajectories, math model architecture, simulation platform, and processes that will be used to evaluate the Contractor-proposed sensors. Four candidate sensors were used in the first phase of the testing. The second phase of testing used four sensors simultaneously: two Marshall Space Flight Center (MSFC) Advanced Video Guidance Sensors (AVGS), a laser-based video sensor that uses retroreflectors attached to the target vehicle, and two commercial laser range finders. The multi-sensor testing was conducted at MSFC's Flight Robotics Laboratory (FRL) using the FRL's 6-DOF gantry system, called the Dynamic Overhead Target System (DOTS). The target vehicle for "docking" in the laboratory was a mockup that was representative of the proposed CEV docking system, with added retroreflectors for the AVGS. The multi-sensor test configuration used 35 open-loop test trajectories covering three major objectives: (1) sensor characterization trajectories designed to test a wide range of performance parameters; (2) CEV-specific trajectories designed to test performance during CEV-like approach and departure profiles; and (3) sensor characterization tests designed for evaluating sensor performance under more extreme conditions as might be induced during a spacecraft failure or during contingency situations. This paper describes the test development, test facility, test preparations, test execution, and test results of the multi-sensor series of trajectories.

A segmented ion engine design for solar electric propulsion systems

NASA Technical Reports Server (NTRS)

Brophy, John R.

1992-01-01

A new ion engine design, called a segmented ion engine, is described which is capable of reducing the required ion source life time for small body rendezvous missions from 18,000 h to about 8,000 h. The use of SAND ion optics for the engine accelerator system makes it possible to substantially reduce the cost of demonstrating the required engine endurance. It is concluded that a flight test of a 5-kW xenon ion propulsion system on the ELITE spacecraft would enormously reduce the cost and risk of using ion propulsion on a planetary vehicle by addressing systems level issues associated with flying a spacecraft radically different from conventional planetary vehicles.

FLASH LIDAR Based Relative Navigation

NASA Technical Reports Server (NTRS)

Brazzel, Jack; Clark, Fred; Milenkovic, Zoran

2014-01-01

Relative navigation remains the most challenging part of spacecraft rendezvous and docking. In recent years, flash LIDARs, have been increasingly selected as the go-to sensors for proximity operations and docking. Flash LIDARS are generally lighter and require less power that scanning Lidars. Flash LIDARs do not have moving parts, and they are capable of tracking multiple targets as well as generating a 3D map of a given target. However, there are some significant drawbacks of Flash Lidars that must be resolved if their use is to be of long-term significance. Overcoming the challenges of Flash LIDARs for navigation-namely, low technology readiness level, lack of historical performance data, target identification, existence of false positives, and performance of vision processing algorithms as intermediaries between the raw sensor data and the Kalman filter-requires a world-class testing facility, such as the Lockheed Martin Space Operations Simulation Center (SOSC). Ground-based testing is a critical step for maturing the next-generation flash LIDAR-based spacecraft relative navigation. This paper will focus on the tests of an integrated relative navigation system conducted at the SOSC in January 2014. The intent of the tests was to characterize and then improve the performance of relative navigation, while addressing many of the flash LIDAR challenges mentioned above. A section on navigation performance and future recommendation completes the discussion.

Data Retrieved by ARCADE-R2 Experiment On Board the BEXUS-17 Balloon

NASA Astrophysics Data System (ADS)

Barbetta, M.; Branz, F.; Carron, A.; Olivieri, L.; Prendin, J.; Sansone, F.; Savioli, L.; Spinello, F.; Francesconi, A.

2015-09-01

The Autonomous Rendezvous, Control And Docking Experiment — Reflight 2 (ARCADE-R2) is a technology demonstrator aiming to prove automatic attitude determination and control, rendezvous and docking capabilities for small scale spacecraft and aircraft. The development of such capabilities could be fundamental to create, in the near future, fleets of cooperative, autonomous unmanned aerial vehicles for mapping, surveillance, inspection and remote observation of hazardous environments; small-class satellites could also benefit from the employment of docking systems to extend and reconfigure their mission profiles. ARCADE-R2 is designed to test these technologies on a stratospheric flight on board the BEXUS-17 balloon, allowing to demonstrate them in a harsh environment subjected to gusty winds and high pressure and temperature variations. In this paper, ARCADE-R2 architecture is introduced and the main results obtained from a stratospheric balloon flight are presented.

ART CONCEPTS - ASTP

NASA Image and Video Library

1975-04-01

S75-27289 (May 1975) --- An artist?s concept depicting the American Apollo spacecraft docked with a Soviet Soyuz spacecraft in Earth orbit. During the joint U.S.-USSR Apollo-Soyuz Test Project mission, scheduled for July 1975, the American and Soviet crews will visit one another?s spacecraft while the Soyuz and Apollo are docked for a maximum period of two days. The mission is designed to test equipment and techniques that will establish international crew rescue capability in space, as well as permit future cooperative scientific missions. Each nation has developed separately docking systems based on a mutually agreeable single set of interface design specifications. The major new U.S. program elements are the docking module and docking system necessary to achieve compatibility of rendezvous and docking systems with the USSR-developed hardware to be used on the Soyuz spacecraft. The DM and docking system together with an Apollo Command/Service Module will be launched by a Saturn 1B launch vehicle. This artwork is by Paul Fjeld.

3D Lasers Increase Efficiency, Safety of Moving Machines

NASA Technical Reports Server (NTRS)

2015-01-01

Canadian company Neptec Design Group Ltd. developed its Laser Camera System, used by shuttles to render 3D maps of their hulls for assessing potential damage. Using NASA funding, the firm incorporated LiDAR technology and created the TriDAR 3D sensor. Its commercial arm, Neptec Technologies Corp., has sold the technology to Orbital Sciences, which uses it to guide its Cygnus spacecraft during rendezvous and dock operations at the International Space Station.

An autonomous rendezvous and docking system using cruise missile technology

NASA Technical Reports Server (NTRS)

Jones, ED; Nicholson, Bruce

1991-01-01

In November 1990 General Dynamics demonstrated an AR&D system for members of the Strategic Avionics Technology Working Group. This simulation utilized prototype hardware derived from the Cruise Missile and Centaur avionics systems. The object of this proof of concept demonstration was to show that all the accuracy, reliability, and operational requirements established for a spacecraft to dock with Space Station Freedom could be met by the proposed AR&D system.

SPHERES as Formation Flight Algorithm Development and Validation Testbed: Current Progress and Beyond

NASA Technical Reports Server (NTRS)

Kong, Edmund M.; Saenz-Otero, Alvar; Nolet, Simon; Berkovitz, Dustin S.; Miller, David W.; Sell, Steve W.

2004-01-01

The MIT-SSL SPHERES testbed provides a facility for the development of algorithms necessary for the success of Distributed Satellite Systems (DSS). The initial development contemplated formation flight and docking control algorithms; SPHERES now supports the study of metrology, control, autonomy, artificial intelligence, and communications algorithms and their effects on DSS projects. To support this wide range of topics, the SPHERES design contemplated the need to support multiple researchers, as echoed from both the hardware and software designs. The SPHERES operational plan further facilitates the development of algorithms by multiple researchers, while the operational locations incrementally increase the ability of the tests to operate in a representative environment. In this paper, an overview of the SPHERES testbed is first presented. The SPHERES testbed serves as a model of the design philosophies that allow for the various researches being carried out on such a facility. The implementation of these philosophies are further highlighted in the three different programs that are currently scheduled for testing onboard the International Space Station (ISS) and three that are proposed for a re-flight mission: Mass Property Identification, Autonomous Rendezvous and Docking, TPF Multiple Spacecraft Formation Flight in the first flight and Precision Optical Pointing, Tethered Formation Flight and Mars Orbit Sample Retrieval for the re-flight mission.

Time-fixed rendezvous by impulse factoring with an intermediate timing constraint. [for transfer orbits

NASA Technical Reports Server (NTRS)

Green, R. N.; Kibler, J. F.; Young, G. R.

1974-01-01

A method is presented for factoring a two-impulse orbital transfer into a three- or four-impulse transfer which solves the rendezvous problem and satisfies an intermediate timing constraint. Both the time of rendezvous and the intermediate time of a alinement are formulated as any element of a finite sequence of times. These times are integer multiples of a constant plus an additive constant. The rendezvous condition is an equality constraint, whereas the intermediate alinement is an inequality constraint. The two timing constraints are satisfied by factoring the impulses into collinear parts that vectorially sum to the original impulse and by varying the resultant period differences and the number of revolutions in each orbit. Five different types of solutions arise by considering factoring either or both of the two impulses into two or three parts with a limit for four total impulses. The impulse-factoring technique may be applied to any two-impulse transfer which has distinct orbital periods.

Preliminary GN&C Design for the On-Orbit Autonomous Assembly of Nanosatellite Demonstration Mission

NASA Technical Reports Server (NTRS)

Pei, Jing; Walsh, Matt; Roithmayr, Carlos; Karlgaard, Chris; Peck, Mason; Murchison, Luke

2017-01-01

Small spacecraft autonomous rendezvous and docking (ARD) is an essential technology for future space structure assembly missions. The On-orbit Autonomous Assembly of Nanosatellites (OAAN) team at NASA Langley Research Center (LaRC) intends to demonstrate the technology to autonomously dock two nanosatellites to form an integrated system. The team has developed a novel magnetic capture and latching mechanism that allows for docking of two CubeSats without precise sensors and actuators. The proposed magnetic docking hardware not only provides the means to latch the CubeSats, but it also significantly increases the likelihood of successful docking in the presence of relative attitude and position errors. The simplicity of the design allows it to be implemented on many CubeSat rendezvous missions. Prior to demonstrating the docking subsystem capabilities on orbit, the GN&C subsystem should have a robust design such that it is capable of bringing the CubeSats from an arbitrary initial separation distance of as many as a few thousand kilometers down to a few meters. The main OAAN Mission can be separated into the following phases: 1) Launch, checkout, and drift, 2) Far-Field Rendezvous or Drift Recovery, 3) Proximity Operations, 4) Docking. This paper discusses the preliminary GN&C design and simulation results for each phase of the mission.

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

NASA Technical Reports Server (NTRS)

Cepollina, Frank J. (Inventor); Corbo, James E. (Inventor); Burns, Richard D. (Inventor); Jedhrich, Nicholas M. (Inventor); Holz, Jill M. (Inventor)

2009-01-01

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR, and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time as well as the ability to intervene using manual override to teleoperate the robot.

Reachability Analysis Applied to Space Situational Awareness

NASA Astrophysics Data System (ADS)

Holzinger, M.; Scheeres, D.

Several existing and emerging applications of Space Situational Awareness (SSA) relate directly to spacecraft Rendezvous, Proximity Operations, and Docking (RPOD) and Formation / Cluster Flight (FCF). When multiple Resident Space Ob jects (RSOs) are in vicinity of one another with appreciable periods between observations, correlating new RSO tracks to previously known objects becomes a non-trivial problem. A particularly difficult sub-problem is seen when long breaks in observations are coupled with continuous, low- thrust maneuvers. Reachability theory, directly related to optimal control theory, can compute contiguous reachability sets for known or estimated control authority and can support such RSO search and correlation efforts in both ground and on-board settings. Reachability analysis can also directly estimate the minimum control authority of a given RSO. For RPOD and FCF applications, emerging mission concepts such as fractionation drastically increase system complexity of on-board autonomous fault management systems. Reachability theory, as applied to SSA in RPOD and FCF applications, can involve correlation of nearby RSO observations, control authority estimation, and sensor track re-acquisition. Additional uses of reachability analysis are formation reconfiguration, worst-case passive safety, and propulsion failure modes such as a "stuck" thruster. Existing reachability theory is applied to RPOD and FCF regimes. An optimal control policy is developed to maximize the reachability set and optimal control law discontinuities (switching) are examined. The Clohessy-Wiltshire linearized equations of motion are normalized to accentuate relative control authority for spacecraft propulsion systems at both Low Earth Orbit (LEO) and Geostationary Earth Orbit (GEO). Several examples with traditional and low thrust propulsion systems in LEO and GEO are explored to illustrate the effects of relative control authority on the time-varying reachability set surface. Both monopropellant spacecraft at LEO and Hall thruster spacecraft at GEO are shown to be strongly actuated while Hall thruster spacecraft at LEO are found to be weakly actuated. Weaknesses with the current implementation are discussed and future numerical improvements and analytical efforts are discussed.

On the relationship between gas and dust in 15 comets: an application to Comet 103P/Hartley 2 target of the NASA EPOXI mission of opportunity

NASA Astrophysics Data System (ADS)

Sanzovo, G. C.; Sanzovo, D. Trevisan; de Almeida, A. A.

After the success of Deep Impact mission to hit the nucleus of Comet 9P/Tempel 1 with an impactor, the concerns are turned now to the possible reutilization of this dormant flyby spacecraft in the study of another comet, for only about 10% of the cost of the original mission. Comet 103P/Hartley 2 on UT 2010 October 11 is the most attractive target in terms of available fuel at rendezvous and arrival time at the comet. In addition, the comet has a low inclination so that major orbital plane changes in the spacecraft trajectory are unnecessary. In an effort to provide information concerning the planning of this new NASA EPOXI space mission of opportunity, we use in this work, visual magnitudes measurements available from International Comet Quarterly (ICQ) to obtain, applying the Semi-Empirical Method of Visual Magnitudes - SEMVM (de Almeida, Singh, & Huebner 1997), the water production rates (in molecules/s) related to its perihelion passage of 1997. When associated to the water vaporization theory of Delsemme (1982), these rates allowed the acquisition of the minimum dimension for the effective nuclear radius of the comet. The water production rates were then converted into gas production rates (in g/s) so that, with the help of the strong correlation between gas and dust found for 12 periodic comets and 3 non-period comets (Trevisan Sanzovo 2006), we obtained the dust loss rates (in g/s), its behavior with the heliocentric distance and the dust-to-gas ratios in this physically attractive rendezvous target-comet to Deep Impact spacecraft at a closest approach of 700 km.

Amor: Investigating The Triple Asteroid System 2001 SN263

NASA Astrophysics Data System (ADS)

Jones, T.; Bellerose, Julie; Lee, P.; Prettyman, T.; Lawrence, D.; Smith, P.; Gaffey, M.; Nolan, M.; Goldsten, J.; Thomas, P.; Veverka, J.; Farquhar, R.; Heldmann, J.; Reddy, V.; Williams, B.; Chartres, J.; DeRosee, R.; Dunham, D.

2010-10-01

The Amor mission will rendezvous and land at the triple Near-Earth Asteroid system (153591) 2001 SN263 and execute detailed, in-situ science investigations. The spacecraft reaches 2001 SN263 by using a two-year ΔVEGA (ΔV-Earth Gravity Assist) trajectory with a relatively low launch C3 of 33.5 km2/s2. Rendezvous will enable reconnaissance activities including global and regional imaging, shape modeling, system dynamics, and compositional mapping. After landing, Amor will conduct in-situ imaging (panoramic to microscopic scale) and compositional measurements to include elemental abundance. The main objectives are to 1) establish in-situ the long-hypothesized link between C-type asteroids and the primitive carbonaceous chondrite (CC) meteorites, 2) investigate the nature, origin and evolution of C-type asteroids, and 3) investigate the origin and evolution of a multiple asteroid system. The mission also addresses the distribution of volatiles and organic materials, impact hazards, and resources for future exploration. Amor is managed by NASA Ames Research Center in partnership with Orbital Sciences, KinetX, MDA, and Draper with heritage instruments provided by Ball Aerospace, JHU/APL, and Firestar Engineering. The science team brings experience from NEAR, Hayabusa, Deep Impact, Dawn, LCROSS, Kepler, and Mars missions. In this paper, we describe the science, mission design, and main operational challenges of performing in-situ science at this triple asteroid system. Challenges include landing on the asteroid components, thermal environment, short day-night cycles, and the operation of deployed instruments in a low gravity (10^-5 g) environment.

Nuclear electric propulsion mission engineering study. Volume 2: Final report

NASA Technical Reports Server (NTRS)

1973-01-01

Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed, along with the impact of its availability on future space programs. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied.

Artist's concept of ASTP mission profile

NASA Image and Video Library

1974-10-01

S74-14949 (October 1974) --- Artist?s drawings and call-outs depict phases of the joint U.S.-USSR Apollo-Soyuz Test Project, an Earth-orbital mission which will feature rendezvous and docking of the respective spacecraft of the two nations. ASTP crewmen for the USSR include Aleksey A. Leonov and Valeriy N. Kubasov. The astronaut team includes astronauts Thomas P. Stafford, Vance D. Brand and Donald K. Slayton. The mission is scheduled to take place in summer 1975.

Earth Observations taken by STS-127 Crew

NASA Image and Video Library

2009-07-30

S127-E-012774 (30 July 2009) --- Backdropped by Earth?s horizon and the blackness of space, a Dual RF Astrodynamic GPS Orbital Navigator Satellite (DRAGONSat) is photographed after its release from Space Shuttle Endeavour?s payload bay by STS-127 crew members. DRAGONSat will look at independent rendezvous of spacecraft in orbit using Global Positioning Satellite data. The two satellites were designed and built by students at the University of Texas, Austin, and Texas A&M University, College Station.

Earth Observations taken by STS-127 Crew

NASA Image and Video Library

2009-07-30

S127-E-012776 (30 July 2009) --- Backdropped by Earth?s horizon and the blackness of space, a Dual RF Astrodynamic GPS Orbital Navigator Satellite (DRAGONSat) is photographed after its release from Space Shuttle Endeavour?s payload bay by STS-127 crew members. DRAGONSat will look at independent rendezvous of spacecraft in orbit using Global Positioning Satellite data. The two satellites were designed and built by students at the University of Texas, Austin, and Texas A&M University, College Station.

Application of a Novel Long-Reach Manipulator Concept to Asteroid Redirect Missions

NASA Technical Reports Server (NTRS)

Dorsey, John T.; Doggett, William R.; Jones, Thomas C.; King, Bruce D.

2015-01-01

A high priority mission currently being formulated by NASA is to capture all or part of an asteroid and return it to cis-lunar space for examination by an astronaut crew. Two major mission architectures are currently being considered: in the first (Mission Concept A), a spacecraft would rendezvous and capture an entire free flying asteroid (up to 14 meters in diameter), and in the second (Mission Concept B), a spacecraft would rendezvous with a large asteroid (which could include one of the Martian moons) and retrieve a boulder (up to 4 meters in diameter). A critical element of the mission is the system that will capture the asteroid or boulder material, enclose it and secure it for the return flight. This paper describes the design concepts, concept of operations, structural sizing and masses of capture systems that are based on a new and novel Tendon- Actuated Lightweight In-Space MANipulator (TALISMAN) general-purpose robotic system. Features of the TALISMAN system are described and the status of its technology development is summarized. TALISMAN-based asteroid material retrieval system concepts and concepts-of-operations are defined for each asteroid mission architecture. The TALISMAN-based capture systems are shown to dramatically increase operational versatility while reducing mission risk. Total masses of TALISMAN-based systems are presented, reinforcing the mission viability of using a manipulator-based approach for the asteroid redirect mission.

Semi-Major Axis Knowledge and GPS Orbit Determination

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell; Schiesser, Emil R.; Bauer, F. (Technical Monitor)

2000-01-01

In recent years spacecraft designers have increasingly sought to use onboard Global Positioning System receivers for orbit determination. The superb positioning accuracy of GPS has tended to focus more attention on the system's capability to determine the spacecraft's location at a particular epoch than on accurate orbit determination, per se. The determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. It is necessary to address semi-major axis mission requirements and the GPS receiver capability for orbital maneuver targeting and other operations that require trajectory prediction. Failure to determine semi-major axis accurately can result in a solution that may not be usable for targeting the execution of orbit adjustment and rendezvous maneuvers. Simple formulas, charts, and rules of thumb relating position, velocity, and semi-major axis are useful in design and analysis of GPS receivers for near circular orbit operations, including rendezvous and formation flying missions. Space Shuttle flights of a number of different GPS receivers, including a mix of unfiltered and filtered solution data and Standard and Precise Positioning Service modes, have been accomplished. These results indicate that semi-major axis is often not determined very accurately, due to a poor velocity solution and a lack of proper filtering to provide good radial and speed error correlation.

Semi-Major Axis Knowledge and GPS Orbit Determination

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell; Schiesser, Emil R.; Bauer, F. (Technical Monitor)

2000-01-01

In recent years spacecraft designers have increasingly sought to use onboard Global Positioning System receivers for orbit determination. The superb positioning accuracy of GPS has tended to focus more attention on the system's capability to determine the spacecraft's location at a particular epoch than on accurate orbit determination, per se. The determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. It is necessary to address semi-major axis mission requirements and the GPS receiver capability for orbital maneuver targeting and other operations that require trajectory prediction. Failure to determine semi-major axis accurately can result in a solution that may not be usable for targeting the execution of orbit adjustment and rendezvous maneuvers. Simple formulas, charts, and rules of thumb relating position, velocity, and semi-major axis are useful in design and analysis of GPS receivers for near circular orbit operations, including rendezvous and formation flying missions. Space Shuttle flights of a number of different GPS receivers, including a mix of unfiltered and filtered solution data and Standard and Precise Positioning, Service modes, have been accomplished. These results indicate that semi-major axis is often not determined very accurately, due to a poor velocity solution and a lack of proper filtering to provide good radial and speed error correlation.

Technology Development of Automated Rendezvous and Docking/Capture Sensors and Docking Mechanism for the Asteroid Redirect Crewed Mission

NASA Technical Reports Server (NTRS)

Hinkel, Heather; Strube, Matthew; Zipay, John J.; Cryan, Scott

2016-01-01

This paper will describe the technology development efforts NASA has underway for Automated Rendezvous and Docking/Capture (AR&D/C) sensors and a docking mechanism and the challenges involved. The paper will additionally address how these technologies will be extended to other missions requiring AR&D/C whether robotic or manned. NASA needs AR&D/C sensors for both the robotic and crewed segments of the Asteroid Redirect Mission (ARM). NASA recently conducted a commonality assessment of the concept of operations for the robotic Asteroid Redirect Vehicle (ARV) and the crewed mission segment using the Orion spacecraft. The commonality assessment also considered several future exploration and science missions requiring an AR&D/C capability. Missions considered were asteroid sample return, satellite servicing, and planetary entry, descent, and landing. This assessment determined that a common sensor suite consisting of one or more visible wavelength cameras, a three-dimensional LIDAR along with long-wavelength infrared cameras for robustness and situational awareness could be used on each mission to eliminate the cost of multiple sensor developments and qualifications. By choosing sensor parameters at build-time instead of at design-time and, without having to requalify flight hardware, a specific mission can design overlapping bearing, range, relative attitude, and position measurement availability to suit their mission requirements with minimal non-recurring engineering costs. The resulting common sensor specification provides the union of all performance requirements for each mission and represents an improvement over the current systems used for AR&D/C today. These sensor specifications are tightly coupled to the docking system capabilities and requirements for final docking conditions. The paper will describe NASA's efforts to develop a standard docking system for use across NASA human spaceflight missions to multiple destinations. It will describe the current design status and the considerations and technologies involved in developing this docking mechanism.

Performance comparison of earth and space storable bipropellant systems in interplanetary missions

NASA Technical Reports Server (NTRS)

Meissinger, H. F.

1978-01-01

The paper evaluates and compares the performance of earth-storable and space-storable liquid bipropellant propulsion systems in high-energy planetary mission applications, including specifically Saturn and Mercury orbiters, as well as asteroid and comet rendezvous missions. The discussion covers a brief review of the status of space-storable propulsion technology, along with an illustrative propulsion module design for a three-axis stabilized outer planet and cometary mission spacecraft of the Mariner class. The results take revised Shuttle/Upper Stage performance projections into account. It is shown that in some of the missions the performance improvement achievable in the ballistic transfer mode with space-storable spacecraft propulsion can provide a possible alternative to the use of solar-electric propulsion.

Mission Steering Profiles of Outer Planetary Orbiters Using Radioisotope Electric Propulsion

NASA Technical Reports Server (NTRS)

Fiehler, Douglas; Oleson, Steven

2004-01-01

Radioisotope Electric Propulsion (REP) has the potential to enable small spacecraft to orbit outer planetary targets with trip times comparable to flyby missions. The ability to transition from a flyby to an orbiter mission lies in the availability of continuous low power electric propulsion along the entire trajectory. The electric propulsion system s role is to add and remove energy from the spacecraft s trajectory to bring it in and out of a heliocentric hyperbolic escape trajectory for the outermost target bodies. Energy is added and the trajectory is reshaped to rendezvous with the closer-in target bodies. Sample REP trajectories will be presented for missions ranging for distances from Jupiter orbit to the Pluto-Kuiper Belt.

An Abstract Plan Preparation Language

NASA Technical Reports Server (NTRS)

Butler, Ricky W.; Munoz, Cesar A.

2006-01-01

This paper presents a new planning language that is more abstract than most existing planning languages such as the Planning Domain Definition Language (PDDL) or the New Domain Description Language (NDDL). The goal of this language is to simplify the formal analysis and specification of planning problems that are intended for safety-critical applications such as power management or automated rendezvous in future manned spacecraft. The new language has been named the Abstract Plan Preparation Language (APPL). A translator from APPL to NDDL has been developed in support of the Spacecraft Autonomy for Vehicles and Habitats Project (SAVH) sponsored by the Explorations Technology Development Program, which is seeking to mature autonomy technology for application to the new Crew Exploration Vehicle (CEV) that will replace the Space Shuttle.

Modelling of 67P cometary grains dynamic in the vicinity of the Rosetta spacecraft

NASA Astrophysics Data System (ADS)

Cipriani, F.; Altobelli, N.; Taylor, M.; Fulle, M.; Della Corte, V.; Rotundi, A.

2017-09-01

The interpretation of a number of Rosetta datasets (e.g. GIADA, COSIMA, MIDAS...), relies on the description of cometary grains dynamic in the close vicinity of the spacecraft. In particular the charged grains behaviour in the 3D spacecraft sheath open to the instrument entrances is complex and has not been described at such scales. The existence of a warm electrons population (a few 10eVs energy) in the cometary plasma as revealed during the Rendez-vous phase has been driving the spacecraft potential to negative values typically in the range -1 to -20V as inferred from RPC measurements [1]. Observation of cometary grains in the 10μm to mm range by GIADA and COSIMA[2] allowed to distinguish so called 'compact' grains of processed materials from the solar nebula from 'fluffy' aggregates of more primitive origin. When detected such grains have been observed to reach the instruments at m/s or less velocities. On particular it was inferred that fluffy aggregates are disrupted by electrostatic forces in the vicinity of the spacecraft due to the effects of local plasma hence resulting in particle showers observed by the instruments.

New surprises in the largest magnetosphere of our solar system.

PubMed

Krupp, Norbert

2007-10-12

En route to its ultimate rendezvous with Pluto, the New Horizons spacecraft passed through the magnetic and plasma environment of Jupiter in February 2007. Onboard instruments collected high-resolution images, spectroscopic data, and information about charged particles. The results have revealed unusual structure and variation in Jupiter's plasma and large plasmoids that travel down the magnetotail. Data on Jupiter's aurora provide details of the interaction with the solar wind, and a major volcanic eruption from the moon Io was observed during the encounter.

SSTAC/ARTS review of the draft Integrated Technology Plan (ITP). Volume 6: Controls and guidance

NASA Technical Reports Server (NTRS)

1991-01-01

Viewgraphs of briefings from the Space Systems and Technology Advisory Committee (SSTAC)/ARTS review of the draft Integrated Technology Plan (ITP) on controls and guidance are included. Topics covered include: strategic avionics technology planning and bridging programs; avionics technology plan; vehicle health management; spacecraft guidance research; autonomous rendezvous and docking; autonomous landing; computational control; fiberoptic rotation sensors; precision instrument and telescope pointing; microsensors and microinstruments; micro guidance and control initiative; and earth-orbiting platforms controls-structures interaction.

IUS/TUG orbital operations and mission support study. Volume 4: Project planning data

NASA Technical Reports Server (NTRS)

1975-01-01

Planning data are presented for the development phases of interim upper stage (IUS) and tug systems. Major project planning requirements, major event schedules, milestones, system development and operations process networks, and relevant support research and technology requirements are included. Topics discussed include: IUS flight software; tug flight software; IUS/tug ground control center facilities, personnel, data systems, software, and equipment; IUS mission events; tug mission events; tug/spacecraft rendezvous and docking; tug/orbiter operations interface, and IUS/orbiter operations interface.

Development of a Deployable Nonmetallic Boom for Reconfigurable Systems of Small Modular Spacecraft

NASA Technical Reports Server (NTRS)

Rehnmark, Fredrik

2007-01-01

Launch vehicle payload capacity and the launch environment represent two of the most operationally limiting constraints on space system mass, volume, and configuration. Large-scale space science and power platforms as well as transit vehicles have been proposed that greatly exceed single-launch capabilities. Reconfigurable systems launched as multiple small modular spacecraft with the ability to rendezvous, approach, mate, and conduct coordinated operations have the potential to make these designs feasible. A key characteristic of these proposed systems is their ability to assemble into desired geometric (spatial) configurations. While flexible and sparse formations may be realized by groups of spacecraft flying in close proximity, flyers physically connected by active structural elements could continuously exchange power, fluids, and heat (via fluids). Configurations of small modular spacecraft temporarily linked together could be sustained as long as needed with minimal propellant use and reconfigured as often as needed over extended missions with changing requirements. For example, these vehicles could operate in extremely compact configurations during boost phases of a mission and then redeploy to generate power or communicate while coasting and upon reaching orbit. In 2005, NASA funded Phase 1 of a program called Modular Reconfigurable High-Energy Technology Demonstrator Assembly Testbed (MRHE) to investigate reconfigurable systems of small spacecraft. The MRHE team was led by NASA's Marshall Space Flight Center and included Lockheed Martin's Advanced Technology Center (ATC) in Palo Alto and its subcontractor, ATK. One of the goals of Phase 1 was to develop an MRHE concept demonstration in a relevant 1-g environment to highlight a number of requisite technologies. In Phase 1 of the MRHE program, Lockheed Martin devised and conducted an automated space system assembly demonstration featuring multipurpose free-floating robots representing Spacecraft in the newly built Controls and Automation Laboratory (CAL) at the ATC. The CAL lab features a 12' x 24' granite air-bearing table and an overhead simulated starfield. Among the technologies needed for the concept demo were mating interfaces allowing the spacecraft to dock and deployable structures allowing for adjustable separation between spacecraft after a rigid connection had been established. The decision to use a nonmetallic deployable boom for this purpose was driven by the MRHE concept demo requirements reproduced in Table 1.

Apollo Soyuz, mission evaluation report

NASA Technical Reports Server (NTRS)

1975-01-01

The Apollo Soyuz mission was the first manned space flight to be conducted jointly by two nations - the United States and the Union of Soviet Socialist Republics. The primary purpose of the mission was to test systems for rendezvous and docking of manned spacecraft that would be suitable for use as a standard international system, and to demonstrate crew transfer between spacecraft. The secondary purpose was to conduct a program of scientific and applications experimentation. With minor modifications, the Apollo and Soyuz spacecraft were like those flown on previous missions. However, a new module was built specifically for this mission - the docking module. It served as an airlock for crew transfer and as a structural base for the docking mechanism that interfaced with a similar mechanism on the Soyuz orbital module. The postflight evaluation of the performance of the docking system and docking module, as well as the overall performance of the Apollo spacecraft and experiments is presented. In addition, the mission is evaluated from the viewpoints of the flight crew, ground support operations, and biomedical operations. Descriptions of the docking mechanism, docking module, crew equipment and experiment hardware are given.

Adaptive nonlinear robust relative pose control of spacecraft autonomous rendezvous and proximity operations.

PubMed

Sun, Liang; Huo, Wei; Jiao, Zongxia

2017-03-01

This paper studies relative pose control for a rigid spacecraft with parametric uncertainties approaching to an unknown tumbling target in disturbed space environment. State feedback controllers for relative translation and relative rotation are designed in an adaptive nonlinear robust control framework. The element-wise and norm-wise adaptive laws are utilized to compensate the parametric uncertainties of chaser and target spacecraft, respectively. External disturbances acting on two spacecraft are treated as a lumped and bounded perturbation input for system. To achieve the prescribed disturbance attenuation performance index, feedback gains of controllers are designed by solving linear matrix inequality problems so that lumped disturbance attenuation with respect to the controlled output is ensured in the L 2 -gain sense. Moreover, in the absence of lumped disturbance input, asymptotical convergence of relative pose are proved by using the Lyapunov method. Numerical simulations are performed to show that position tracking and attitude synchronization are accomplished in spite of the presence of couplings and uncertainties. Copyright © 2016 ISA. Published by Elsevier Ltd. All rights reserved.

State dependent model predictive control for orbital rendezvous using pulse-width pulse-frequency modulated thrusters

NASA Astrophysics Data System (ADS)

Li, Peng; Zhu, Zheng H.; Meguid, S. A.

2016-07-01

This paper studies the pulse-width pulse-frequency modulation based trajectory planning for orbital rendezvous and proximity maneuvering near a non-cooperative spacecraft in an elliptical orbit. The problem is formulated by converting the continuous control input, output from the state dependent model predictive control, into a sequence of pulses of constant magnitude by controlling firing frequency and duration of constant-magnitude thrusters. The state dependent model predictive control is derived by minimizing the control error of states and control roughness of control input for a safe, smooth and fuel efficient approaching trajectory. The resulting nonlinear programming problem is converted into a series of quadratic programming problem and solved by numerical iteration using the receding horizon strategy. The numerical results show that the proposed state dependent model predictive control with the pulse-width pulse-frequency modulation is able to effectively generate optimized trajectories using equivalent control pulses for the proximity maneuvering with less energy consumption.

Official Portrait of Astronaut Michael Collins

NASA Technical Reports Server (NTRS)

1967-01-01

This is the official NASA portrait of astronaut Michael Collins. Collins chose an Air Force career following graduation from West Point. He served as an experimental flight test officer at the Air Force Flight Test Center, Edwards Air Force Base, California, and, in that capacity, tested performance and stability and control characteristics of Air Force aircraft, primarily jet fighters. Having logged approximately 5,000 hours flying time, Collins was one of the third group of astronauts named by NASA in October 1963. Collins completed two space flights, logging 266 hours in space, of which, 1 hour and 27 minutes was spent in Extra Vehicular Activity (EVA). On July 18, 1966, he served as backup pilot for the Gemini VII mission which included a successful rendezvous and docking with a separately launched Agena target vehicle and, using the power of the Agena, maneuvered the Gemini spacecraft into another orbit for a rendezvous with a second, passive Agena. His skillful performance in completing two periods of EVA included the recovery of a micrometeorite detection experiment from the passive Agena. July 16-24, 1969, Collins served as command module (CM) pilot on Apollo 11, the historic first lunar landing mission. He remained aboard the CM, Columbia, on station in lunar orbit and performed the final re-docking maneuvers following a successful lunar orbit rendezvous with the Lunar Module (LM), Eagle. Collins left NASA in January 1970.

Saturn Apollo Program

NASA Image and Video Library

1967-01-09

This is the official NASA portrait of astronaut Michael Collins. Collins chose an Air Force career following graduation from West Point. He served as an experimental flight test officer at the Air Force Flight Test Center, Edwards Air Force Base, California, and, in that capacity, tested performance and stability and control characteristics of Air Force aircraft, primarily jet fighters. Having logged approximately 5,000 hours flying time, Collins was one of the third group of astronauts named by NASA in October 1963. Collins completed two space flights, logging 266 hours in space, of which, 1 hour and 27 minutes was spent in Extra Vehicular Activity (EVA). On July 18, 1966, he served as backup pilot for the Gemini VII mission which included a successful rendezvous and docking with a separately launched Agena target vehicle and, using the power of the Agena, maneuvered the Gemini spacecraft into another orbit for a rendezvous with a second, passive Agena. His skillful performance in completing two periods of EVA included the recovery of a micrometeorite detection experiment from the passive Agena. July 16-24, 1969, Collins served as command module (CM) pilot on Apollo 11, the historic first lunar landing mission. He remained aboard the CM, Columbia, on station in lunar orbit and performed the final re-docking maneuvers following a successful lunar orbit rendezvous with the Lunar Module (LM), Eagle. Collins left NASA in January 1970.

Solar Electric Propulsion for Primitive Body Science Missions

NASA Technical Reports Server (NTRS)

Witzberger, Kevin E.

2006-01-01

This paper describes work that assesses the performance of solar electric propulsion (SEP) for three different primitive body science missions: 1) Comet Rendezvous 2) Comet Surface Sample Return (CSSR), and 3) a Trojan asteroid/Centaur object Reconnaissance Flyby. Each of these missions launches from Earth between 2010 and 2016. Beginning-of-life (BOL) solar array power (referenced at 1 A.U.) varies from 10 to 18 kW. Launch vehicle selections range from a Delta II to a Delta IV medium-class. The primary figure of merit (FOM) is net delivered mass (NDM). This analysis considers the effects of imposing various mission constraints on the Comet Rendezvous and CSSR missions. Specifically, the Comet Rendezvous mission analysis examines an arrival date constraint with a launch year variation, whereas the CSSR mission analysis investigates an Earth entry velocity constraint commensurate with past and current missions. Additionally, the CSSR mission analysis establishes NASA's New Frontiers (NF) Design Reference Mission (DRM) in order to evaluate current and future SEP technologies. The results show that transfer times range from 5 to 9 years (depending on the mission). More importantly, the spacecraft's primary propulsion system performs an average 5-degree plane change on the return leg of the CSSR mission to meet the previously mentioned Earth entry velocity constraint. Consequently, these analyses show that SEP technologies that have higher thrust-to-power ratios can: 1) reduce flight time, and 2) change planes more efficiently.

The solar panels of the spacecraft Stardust are deployed before undergoing lighting test in the PHSF

NASA Technical Reports Server (NTRS)

1999-01-01

In the Payload Hazardous Servicing Facility, workers look over the solar panels on the Stardust spacecraft that are deployed for lighting tests. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule to be jettisoned as it swings by Earth in January 2006.

Spacecraft Autonomy and Automation: A Comparative Analysis of Strategies for Cost Effective Mission Operations

NASA Technical Reports Server (NTRS)

Wright, Nathaniel, Jr.

2000-01-01

The evolution of satellite operations over the last 40 years has drastically changed. October 4, 1957 (during the cold war) the Soviet Union launched the world's first spacecraft into orbit. The Sputnik satellite orbited Earth for three months and catapulted the United States into a race for dominance in space. A year after Sputnik, President Dwight Eisenhower formed the National Space and Aeronautics Administration (NASA). With a team of scientists and engineers, NASA successfully launched Explorer 1, the first US satellite to orbit Earth. During these early years, massive amounts of ground support equipment and operators were required to successfully operate spacecraft vehicles. Today, budget reductions and technological advances have forced new approaches to spacecraft operations. These approaches require increasingly complex, on board spacecraft systems, that enable autonomous operations, resulting in more cost-effective mission operations. NASA's Goddard Space Flight Center, considered world class in satellite development and operations, has developed and operated over 200 satellites during its 40 years of existence. NASA Goddard is adopting several new millennium initiatives that lower operational costs through the spacecraft autonomy and automation. This paper examines NASA's approach to spacecraft autonomy and ground system automation through a comparative analysis of satellite missions for Hubble Space Telescope-HST, Near Earth Asteroid Rendezvous-NEAR, and Solar Heliospheric Observatory-SoHO, with emphasis on cost reduction methods, risk analysis and anomalies and strategies employed for mitigating risk.

The Near Earth Object Scout Spacecraft: A Low Cost Approach to in-situ Characterization of the NEO Population

NASA Technical Reports Server (NTRS)

Koontz, Steven L.; Condon, Gerald; Graham, Lee; Bevilacqua, Ricardo

2014-01-01

In this paper we describe a micro/nano satellite spacecraft and a supporting mission profile and architecture designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonable cost. The spacecraft will be referred to as the NEO Scout. NEO Scout spacecraft are to be placed in GTO, GEO, or cis-lunar space as secondary payloads on launch vehicles headed for GTO or beyond and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO scout system is to design the mission timeline and spacecraft to rendezvous with and land on the target NEOs during close approach to the Earth-Moon system using low-thrust/high- impulse propulsion systems. Mission feasibility and preliminary design analysis are presented along with detailed trajectory calculations. The use of micro/nano satellites in low-cost interplanetary exploration is attracting increasing attention and is the subject of several annual workshops and published design studies (1-4). The NEO population consists of those asteroids and short period comets orbiting the Sun with a perihelion of 1.3 astronomical units or less (5-8). As of July 30, 2013 10065 Near-Earth objects have been discovered. The spin rate, mass, density, surface physical (especially mechanical) properties, composition, and mineralogy of the vast majority of these objects are highly uncertain and the limited available telescopic remote sensing data imply a very diverse population (5-8). In-situ measurements by robotic spacecraft are urgently needed to provide the characterization data needed to support hardware and mission design for more ambitious human and robotic NEO operations. Large numbers of NEOs move into close proximity with the Earth-Moon system every year (9). The JPL Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) (10) has produced detailed mission profile and delta V requirements for various NEO missions ranging from 30 to 420 days in duration and assuming chemical propulsion. Similar studies have been reported assuming high power electric propulsion for manned NEO rendezvous missions (11). The delta V requirement breakdown and mission profile data from references 10 and 11 are used as a basis for sizing the NEO Scout spacecraft and for conducting preliminary feasibility assessments using the Tsiokolvsky rocket equation, a (worst-case) delta V requirement of 10 km/sec, and a maximum spacecraft dry mass of 20 kg. Using chemical propellant for a 10 km/sec delta V drives spacecraft wet mass well above 300 kg so that chemical propulsion is a non-starter for the proposed mission profile and spacecraft wet mass limits. In contrast, a solar electric propulsion system needs only 8 kg of Xe propellant to accelerate the spacecraft to 10 km/sec in 163 days with 0.02 N of thrust and 500 W of power from1.6 sq m of 29% efficient solar panels. In a second example, accelerating a 4 kg payload to 7 km/sec over 180 days requires about 6.7 kg of propellant and 1.2 kg of solar panels (12 kg total spacecraft wet mass).

Modeling and Simulation of a Novel Relay Node Based Secure Routing Protocol Using Multiple Mobile Sink for Wireless Sensor Networks.

PubMed

Perumal, Madhumathy; Dhandapani, Sivakumar

2015-01-01

Data gathering and optimal path selection for wireless sensor networks (WSN) using existing protocols result in collision. Increase in collision further increases the possibility of packet drop. Thus there is a necessity to eliminate collision during data aggregation. Increasing the efficiency is the need of the hour with maximum security. This paper is an effort to come up with a reliable and energy efficient WSN routing and secure protocol with minimum delay. This technique is named as relay node based secure routing protocol for multiple mobile sink (RSRPMS). This protocol finds the rendezvous point for optimal transmission of data using a "splitting tree" technique in tree-shaped network topology and then to determine all the subsequent positions of a sink the "Biased Random Walk" model is used. In case of an event, the sink gathers the data from all sources, when they are in the sensing range of rendezvous point. Otherwise relay node is selected from its neighbor to transfer packets from rendezvous point to sink. A symmetric key cryptography is used for secure transmission. The proposed relay node based secure routing protocol for multiple mobile sink (RSRPMS) is experimented and simulation results are compared with Intelligent Agent-Based Routing (IAR) protocol to prove that there is increase in the network lifetime compared with other routing protocols.

Design and Implementation of the Automated Rendezvous Targeting Algorithms for Orion

NASA Technical Reports Server (NTRS)

DSouza, Christopher; Weeks, Michael

2010-01-01

The Orion vehicle will be designed to perform several rendezvous missions: rendezvous with the ISS in Low Earth Orbit (LEO), rendezvous with the EDS/Altair in LEO, a contingency rendezvous with the ascent stage of the Altair in Low Lunar Orbit (LLO) and a contingency rendezvous in LLO with the ascent and descent stage in the case of an aborted lunar landing. Therefore, it is not difficult to realize that each of these scenarios imposes different operational, timing, and performance constraints on the GNC system. To this end, a suite of on-board guidance and targeting algorithms have been designed to meet the requirement to perform the rendezvous independent of communications with the ground. This capability is particularly relevant for the lunar missions, some of which may occur on the far side of the moon. This paper will describe these algorithms which are designed to be structured and arranged in such a way so as to be flexible and able to safely perform a wide variety of rendezvous trajectories. The goal of the algorithms is not to merely fly one specific type of canned rendezvous profile. Conversely, it was designed from the start to be general enough such that any type of trajectory profile can be flown.(i.e. a coelliptic profile, a stable orbit rendezvous profile, and a expedited LLO rendezvous profile, etc) all using the same rendezvous suite of algorithms. Each of these profiles makes use of maneuver types which have been designed with dual goals of robustness and performance. They are designed to converge quickly under dispersed conditions and they are designed to perform many of the functions performed on the ground today. The targeting algorithms consist of a phasing maneuver (NC), an altitude adjust maneuver (NH), and plane change maneuver (NPC), a coelliptic maneuver (NSR), a Lambert targeted maneuver, and several multiple-burn targeted maneuvers which combine one of more of these algorithms. The derivation and implementation of each of these algorithms will be discussed in detail, as well and the Rendezvous Targeting "wrapper" which will sequentially tie them all together into a single onboard targeting tool which can produce a final integrated rendezvous trajectory. In a similar fashion, the various guidance modes available for flying out each of these maneuvers will be discussed as well. This paradigm of having the onboard guidance & targeting capability described above is different than the way the Space Shuttle has operated thus far. As a result, a discussion of these differences in terms of operations and ground and crew intervention will also be discussed. However, the general framework of how the mission designers on the ground first perform all mission design and planning functions, and then uplink that burn plan to the vehicle ensures that the ground will be involved to ensure safety and reliability. The only real difference is which of these functions will be done onboard vs. on the ground as done currently. Finally, this paper will describe the performance of each of these algorithms individually as well as the entire suite of algorithms as applied to the Orion ISS and EDS/Altair rendezvous missions in LEO. These algorithms have been incorporated in both a Linear Covariance environment and a Monte Carlo environment and the results of these dispersion analyses will be presented in the paper as well.

Fully autonomous navigation for the NASA cargo transfer vehicle

NASA Technical Reports Server (NTRS)

Wertz, James R.; Skulsky, E. David

1991-01-01

A great deal of attention has been paid to navigation during the close approach (less than or equal to 1 km) phase of spacecraft rendezvous. However, most spacecraft also require a navigation system which provides the necessary accuracy for placing both satellites within the range of the docking sensors. The Microcosm Autonomous Navigation System (MANS) is an on-board system which uses Earth-referenced attitude sensing hardware to provide precision orbit and attitude determination. The system is capable of functioning from LEO to GEO and beyond. Performance depends on the number of available sensors as well as mission geometry; however, extensive simulations have shown that MANS will provide 100 m to 400 m (3(sigma)) position accuracy and 0.03 to 0.07 deg (3(sigma)) attitude accuracy in low Earth orbit. The system is independent of any external source, including GPS. MANS is expected to have a significant impact on ground operations costs, mission definition and design, survivability, and the potential development of very low-cost, fully autonomous spacecraft.

Construction of optimum controls and trajectories of motion of the center of masses of a spacecraft equipped with the solar sail and low-thrust engine, using quaternions and Kustaanheimo-Stiefel variables

NASA Astrophysics Data System (ADS)

Sapunkov, Ya. G.; Chelnokov, Yu. N.

2014-11-01

The problem of optimum rendezvous of a controllable spacecraft (SC) with an uncontrollable spacecraft, moving over a Keplerian elliptic orbit in the gravitational field of the Sun, is considered. Control of the SC is performed using a solar sail and low-thrust engine. For solving the problem, the regular quaternion equations of the two-body problem with the Kustaanheimo-Stiefel variables and the Pontryagin maximum principle are used. The combined integral quality functional, which characterizes energy consumption for controllable SC transition from an initial to final state and the time spent for this transition, is used as a minimized functional. The differential boundary-value optimization problems are formulated, and their first integrals are found. Examples of numerical solution of problems are presented. The paper develops the application [1-6] of quaternion regular equations with the Kustaanheimo-Stiefel variables in the space flight mechanics.

OAST Space Theme Workshop. Volume 3: Working group summary. 1: Navigation, guidance, control (E-1) A. Statement. B. Technology needs (form 1). C. Priority assessment (form 2)

NASA Technical Reports Server (NTRS)

1976-01-01

The six themes identified by the Workshop have many common navigation guidance and control needs. All the earth orbit themes have a strong requirement for attitude, figure and stabilization control of large space structures, a requirement not currently being supported. All but the space transportation theme have need for precision pointing of spacecraft and instruments. In addition all the themes have requirements for increasing autonomous operations for such activities as spacecraft and experiment operations, onboard mission modification, rendezvous and docking, spacecraft assembly and maintenance, navigation and guidance, and self-checkout, test and repair. Major new efforts are required to conceptualize new approaches to large space antennas and arrays that are lightweight, readily deployable, and capable of precise attitude and figure control. Conventional approaches offer little hope of meeting these requirements. Functions that can benefit from increasing automation or autonomous operations are listed.

Dawn Orbit Determination Team: Trajectory and Gravity Prediction Performance During Vesta Science Phases

NASA Technical Reports Server (NTRS)

Kennedy, Brian; Abrahamson, Matt; Ardito, Alessandro; Han, Dongsuk; Haw, Robert; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew

2013-01-01

The Dawn spacecraft was launched on September 27th, 2007. Its mission is to consecutively rendezvous with and observe the two largest bodies in the asteroid belt, Vesta and Ceres. It has already completed over a year's worth of direct observations of Vesta (spanning from early 2011 through late 2012) and is currently on a cruise trajectory to Ceres, where it will begin scientific observations in mid-2015. Achieving this data collection required careful planning and execution from all spacecraft teams. Dawn's Orbit Determination (OD) team was tasked with accurately predicting the trajectory of the Dawn spacecraft during the Vesta science phases, and also determining the parameters of Vesta to support future science orbit design. The future orbits included the upcoming science phase orbits as well as the transfer orbits between science phases. In all, five science phases were executed at Vesta, and this paper will describe some of the OD team contributions to the planning and execution of those phases.

Deep Space 1 Using its Ion Engine (Artist's Concept)

NASA Technical Reports Server (NTRS)

2003-01-01

NASA's New Millennium Deep Space 1 spacecraft approaching the comet 19P/Borrelly. With its primary mission to serve as a technology demonstrator--testing ion propulsion and 11 other advanced technologies--successfully completed in September 1999, Deep Space 1 is now headed for a risky, exciting rendezvous with Comet Borrelly. NASA extended the mission, taking advantage of the ion propulsion and other systems to target the daring encounter with the comet in September 2001. Once a sci-fi dream, the ion propulsion engine has powered the spacecraft for over 12,000 hours. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The first flight in NASA's New Millennium Program, Deep Space 1 was launched October 24, 1998 aboard a Boeing Delta 7326 rocket from Cape Canaveral Air Station, FL. Deep Space 1 successfully completed and exceeded its mission objectives in July 1999 and flew by a near-Earth asteroid, Braille (1992 KD), in September 1999.

Relationship between automation trust and operator performance for the novice and expert in spacecraft rendezvous and docking (RVD).

PubMed

Niu, Jianwei; Geng, He; Zhang, Yijing; Du, Xiaoping

2018-09-01

Operator trust in automation is a crucial factor influencing its use and operational performance. However, the relationship between automation trust and performance remains poorly understood and requires further investigation. The objective of this paper is to explore the difference in trust and performance on automation-aided spacecraft rendezvous and docking (RVD) between the novice and the expert and to investigate the relationship between automation trust and performance as well. We employed a two-factor mixed design, with training skill (novice and expert) and automation mode (manual RVD and automation aided RVD) serving as the two factors. Twenty participants, 10 novices and 10 experts, were recruited to conduct six RVD tasks for two automation levels. After the tasks, operator performance was recorded by the desktop hand-held docking training equipment. Operator trust was also measured by a 12-items questionnaire at the beginning and end of each trial. As a result, automation narrowed the performance gap significantly between the novice and the expert, and the automation trust showed a marginally significant difference between the novice and the expert. Furthermore, the result demonstrated that the attitude angle control error of the expert was related to the total trust score, whereas other automation performance indicators were not related to the total score of trust. However, automation performance was related to the dimensions of trust, such as entrust, harmful, and dependable. Copyright © 2018 Elsevier Ltd. All rights reserved.

Reactive Sequencing for Autonomous Navigation Evolving from Phoenix Entry, Descent, and Landing

NASA Technical Reports Server (NTRS)

Grasso, Christopher A.; Riedel, Joseph E.; Vaughan, Andrew T.

2010-01-01

Virtual Machine Language (VML) is an award-winning advanced procedural sequencing language in use on NASA deep-space missions since 1997, and was used for the successful entry, descent, and landing (EDL) of the Phoenix spacecraft onto the surface of Mars. Phoenix EDL utilized a state-oriented operations architecture which executed within the constraints of the existing VML 2.0 flight capability, compatible with the linear "land or die" nature of the mission. The intricacies of Phoenix EDL included the planned discarding of portions of the vehicle, the complex communications management for relay through on-orbit assets, the presence of temporally indeterminate physical events, and the need to rapidly catch up four days of sequencing should a reboot of the spacecraft flight computer occur shortly before atmospheric entry. These formidable operational challenges led to new techniques for packaging and coordinating reusable sequences called blocks using one-way synchronization via VML sequencing global variable events. The coordinated blocks acted as an ensemble to land the spacecraft, while individually managing various elements in as simple a fashion as possible. This paper outlines prototype VML 2.1 flight capabilities that have evolved from the one-way synchronization techniques in order to implement even more ambitious autonomous mission capabilities. Target missions for these new capabilities include autonomous touch-and-go sampling of cometary and asteroidal bodies, lunar landing of robotic missions, and ultimately landing of crewed lunar vehicles. Close proximity guidance, navigation, and control operations, on-orbit rendezvous, and descent and landing events featured in these missions require elaborate abort capability, manifesting highly non-linear scenarios that are so complex as to overtax traditional sequencing, or even the sort of one-way coordinated sequencing used during EDL. Foreseeing advanced command and control needs for small body and lunar landing guidance, navigation and control scenarios, work began three years ago on substantial upgrades to VML that are now being exercised in scenarios for lunar landing and comet/asteroid rendezvous. The advanced state-based approach includes coordinated state transition machines with distributed decision-making logic. These state machines are not merely sequences - they are reactive logic constructs capable of autonomous decision making within a well-defined domain. Combined with the JPL's AutoNav software used on Deep Space 1 and Deep Impact, the system allows spacecraft to autonomously navigate to an unmapped surface, soft-contact, and either land or ascend. The state machine architecture enabled by VML 2.1 has successfully performed sampling missions and lunar descent missions in a simulated environment, and is progressing toward flight capability. The authors are also investigating using the VML 2.1 flight director architecture to perform autonomous activities like rendezvous with a passive hypothetical Mars sample return capsule. The approach being pursued is similar to the touch-and-go sampling state machines, with the added complications associated with the search for, physical capture of, and securing of a separate spacecraft. Complications include optically finding and tracking the Orbiting Sample Capsule (OSC), keeping the OSC illuminated, making orbital adjustments, and physically capturing the OSC. Other applications could include autonomous science collection and fault compensation.

Hydra Rendezvous and Docking Sensor

NASA Technical Reports Server (NTRS)

Roe, Fred; Carrington, Connie

2007-01-01

The U.S. technology to support a CEV AR&D activity is mature and was developed by NASA and supporting industry during an extensive research and development program conducted during the 1990's and early 2000 time frame at the Marshall Space Flight Center. Development and demonstration of a rendezvous/docking sensor was identified early in the AR&D Program as the critical enabling technology that allows automated proxinity operations and docking. A first generation rendezvous/docking sensor, the Video Guidance Sensor (VGS) was developed and successfully flown on STS 87 and again on STS 95, proving the concept of a video-based sensor. Advances in both video and signal processing technologies and the lessons learned from the two successful flight experiments provided a baseline for the development of a new generation of video based rendezvous/docking sensor. The Advanced Video Guidance Sensor (AVGS) has greatly increased performance and additional capability for longer-range operation. A Demonstration Automatic Rendezvous Technology (DART) flight experiment was flown in April 2005 using AVGS as the primary proximity operations sensor. Because of the absence of a docking mechanism on the target satellite, this mission did not demonstrate the ability of the sensor to coltrold ocking. Mission results indicate that the rendezvous sensor operated successfully in "spot mode" (2 km acquisition of the target, bearing data only) but was never commanded to "acquire and track" the docking target. Parts obsolescence issues prevent the construction of current design AVGS units to support the NASA Exploration initiative. This flight proven AR&D technology is being modularized and upgraded with additional capabilities through the Hydra project at the Marshall Space Flight Center. Hydra brings a unique engineering approach and sensor architecture to the table, to solve the continuing issues of parts obsolescence and multiple sensor integration. This paper presents an approach to sensor hardware trades, to address the needs of future vehicles that may rendezvous and dock with the International Space Station (ISS). It will also discuss approaches for upgrading AVGS to address parts obsolescence, and concepts for modularizing the sensor to provide configuration flexibility for multiple vehicle applications. Options for complementary sensors to be integrated into the multi-head Hydra system will also be presented. Complementary sensor options include ULTOR, a digital image correlator system that could provide relative six-degree-of-freedom information independently from AVGS, and time-of-flight sensors, which determine the range between vehicles by timing pulses that travel from the sensor to the target and back. Common targets and integrated targets, suitable for use with the multi-sensor options in Hydra, will also be addressed.

Proximity Operations Nano-Satellite Flight Demonstration (PONSFD) Rendezvous Proximity Operations Design and Trade Studies

NASA Astrophysics Data System (ADS)

Griesbach, J.; Westphal, J. J.; Roscoe, C.; Hawes, D. R.; Carrico, J. P.

2013-09-01

The Proximity Operations Nano-Satellite Flight Demonstration (PONSFD) program is to demonstrate rendezvous proximity operations (RPO), formation flying, and docking with a pair of 3U CubeSats. The program is sponsored by NASA Ames via the Office of the Chief Technologist (OCT) in support of its Small Spacecraft Technology Program (SSTP). The goal of the mission is to demonstrate complex RPO and docking operations with a pair of low-cost 3U CubeSat satellites using passive navigation sensors. The program encompasses the entire system evolution including system design, acquisition, satellite construction, launch, mission operations, and final disposal. The satellite is scheduled for launch in Fall 2015 with a 1-year mission lifetime. This paper provides a brief mission overview but will then focus on the current design and driving trade study results for the RPO mission specific processor and relevant ground software. The current design involves multiple on-board processors, each specifically tasked with providing mission critical capabilities. These capabilities range from attitude determination and control to image processing. The RPO system processor is responsible for absolute and relative navigation, maneuver planning, attitude commanding, and abort monitoring for mission safety. A low power processor running a Linux operating system has been selected for implementation. Navigation is one of the RPO processor's key tasks. This entails processing data obtained from the on-board GPS unit as well as the on-board imaging sensors. To do this, Kalman filters will be hosted on the processor to ingest and process measurements for maintenance of position and velocity estimates with associated uncertainties. While each satellite carries a GPS unit, it will be used sparsely to conserve power. As such, absolute navigation will mainly consist of propagating past known states, and relative navigation will be considered to be of greater importance. For relative observations, each spacecraft hosts 3 electro-optical sensors dedicated to imaging the companion satellite. The image processor will analyze the images to obtain estimates for range, bearing, and pose, with associated rates and uncertainties. These observations will be fed to the RPO processor's relative Kalman filter to perform relative navigation updates. This paper includes estimates for expected navigation accuracies for both absolute and relative position and velocity. Another key task for the RPO processor is maneuver planning. This includes automation to plan maneuvers to achieve a desired formation configuration or trajectory (including docking), as well as automation to safely react to potentially dangerous situations. This will allow each spacecraft to autonomously plan fuel-efficient maneuvers to achieve a desired trajectory as well as compute adjustment maneuvers to correct for thrusting errors. This paper discusses results from a trade study that has been conducted to examine maneuver targeting algorithms required on-board the spacecraft. Ground software will also work in conjunction with the on-board software to validate and approve maneuvers as necessary.

Rosetta enters hibernation

NASA Astrophysics Data System (ADS)

Ferri, Paolo; Accomazzo, Andrea; Hubault, Armelle; Lodiot, Sylvain; Pellon-Bailon, Jose-Luis; Porta, Roberto

2012-10-01

The International Rosetta Mission was launched on 2nd March 2004 on its 10 years journey to comet 67P/Churyumov-Gerasimenko. Rosetta will reach the comet in 2014, orbit it for about 1.5 years down to distances of a few kilometres and deliver the Lander Philae onto its surface. Following the fly-by of Asteroid (21-)Lutetia in 2010, Rosetta continued its travel towards the planned comet encounter in 2014. In this phase Rosetta became the solar-powered spacecraft that reached the largest Sun distances in history of spaceflight, up to an aphelion at 5.3 AU in October 2012. At distances above 4.5 AU the spacecraft's solar generator power is not sufficient to keep all spacecraft systems active. Therefore in June 2011 the spacecraft was spun up to provide gyroscopic stabilisation, and most of its on-board units, including those used for attitude control and communications, were switched off. Over this "hibernation" phase of about 2.5 years the spacecraft will keep a minimum of autonomy active to ensure maintenance of safe thermal conditions. After Lutetia fly-by, flight controllers had to tackle two anomalies that had significant impacts on the mission operations. A leak in the reaction control subsystem was confirmed and led to the re-definition of the operational strategy to perform the comet rendezvous manoeuvres planned for 2011 and 2014. Anomalous jumps detected in the estimated friction torque of two of the four reaction wheels used for attitude control forced the rapid adoption of measures to slow down the wheels degradation. This included in-flight re-lubrication activities and changes in the wheels operational speed regime. Once the troubleshooting of the two anomalies was completed, and the related operational scenarios were implemented, the first large (790 m/s) comet rendezvous manoeuvre was executed, split into several long burns in January and February 2011. The second burn was unexpectedly interrupted due to the anomalous behaviour of two thrusters, causing attitude off-pointing. Flight controllers modified the thrusters operation parameters in the on-board software and managed to re-start the sequence of burns and successfully complete the manoeuvre. After the manoeuvre, preparation for the critical spin-up and hibernation entry activities, planned for June 2011, began. This paper presents the activities carried out on Rosetta in the final year before hibernation entry. The major anomalies and the related troubleshooting and workaround solutions are detailed. Lessons learned from the operation of the first spacecraft operating with solar power at Jupiter-like distances from the Sun are presented and discussed.

Differential Drag Demonstration: A Post-Mission Experiment with the EO-1 Spacecraft

NASA Technical Reports Server (NTRS)

Hull, Scott; Shelton, Amanda; Richardson, David

2017-01-01

Differential drag is a technique for altering the semi-major axis, velocity, and along-track position of a spacecraft in low Earth orbit. It involves varying the spacecrafts cross-sectional area relative to its velocity direction by temporarily changing attitude and solar array angles, thus varying the amount of atmospheric drag on the spacecraft. The technique has recently been proposed and used by at least three satellite systems for initial separation of constellation spacecraft after launch, stationkeeping during the mission, and potentially for conjunction avoidance. Similarly, differential drag has been proposed as a control strategy for rendezvous, removing the need for active propulsion. In theory, some operational missions that lack propulsion capability could use this approach for conjunction avoidance, though options are typically constrained for spacecraft that are already in orbit. Shortly before the spacecraft was decommissioned, an experiment was performed using NASAs EO-1 spacecraft in order to demonstrate differential drag on an operational spacecraft in orbit, and discover some of the effects differential drag might manifest. EO-1 was not designed to maintain off-nominal orientations for long periods, and as a result the team experienced unanticipated challenges during the experiment. This paper will discuss operations limitations identified before the experiment, as well as those discovered during the experiment. The effective displacement that resulted from increasing the drag area for 39 hours will be compared to predictions as well as the expected position if the spacecraft maintained nominal operations. A hypothetical scenario will also be examined, studying the relative risks of maintaining an operational spacecraft bus in order to maintain the near-maximum drag area orientation and hasten reentry.

Differential Drag Demonstration: A Post-Mission Experiment with the EO-1 Spacecraft

NASA Technical Reports Server (NTRS)

Hull, Scott; Shelton, Amanda; Richardson, David

2017-01-01

Differential drag is a technique for altering the semimajor axis, velocity, and along-track position of a spacecraft in low Earth orbit. It involves varying the spacecraft's cross-sectional area relative to its velocity direction by temporarily changing attitude and solar array angles, thus varying the amount of atmospheric drag on the spacecraft. The technique has recently been proposed and used by at least three satellite systems for initial separation of constellation spacecraft after launch, stationkeeping during the mission, and potentially for conjunction avoidance. Similarly, differential drag has been proposed as a control strategy for rendezvous, removing the need for active propulsion. In theory, some operational missions that lack propulsion capability could use this approach for conjunction avoidance, though options are typically constrained for spacecraft that are already in orbit. Shortly before the spacecraft was decommissioned, an experiment was performed using NASA's EO-1 spacecraft in order to demonstrate differential drag on an operational spacecraft in orbit, and discover some of the effects differential drag might manifest. EO-1 was not designed to maintain off-nominal orientations for long periods, and as a result the team experienced unanticipated challenges during the experiment. This paper will discuss operations limitations identified before the experiment, as well as those discovered during the experiment. The effective displacement that resulted from increasing the drag area for 39 hours will be compared to predictions as well as the expected position if the spacecraft maintained nominal operations. A hypothetical scenario will also be examined, studying the relative risks of maintaining an operational spacecraft bus in order to maintain the near-maximum drag area orientation and hasten reentry.

Comets - Groundbased observations of spacecraft mission candidates

NASA Technical Reports Server (NTRS)

Osip, David J.; Schleicher, David G.; Millis, Robert L.

1992-01-01

Ground-based narrowband photometry results are presented for nine candidate comets for flyby and/or rendezvous missions. The comets include Churyumov-Gerasimenko, d'Arrest, Encke, Grigg-Skjellerup, Honda-Mrkos-Pajdusakova, Kopff, Tempel 1, Tempel 2, and Wild 2. On the basis of measured OH production rates and a model of the sublimation of water from the surface, limits are derived on the size of each cometary nucleus. A detailed analysis of the characteristics of these nine viable mission candidates can furnish the bases for the prioritization of targets of prospective missions.

First Integrated Flight Simulation For STS 114

NASA Image and Video Library

2004-10-13

JSC2004-E-45138 (13 October 2004) --- Astronaut Stephen N. Frick monitors communications at the spacecraft communicator (CAPCOM) console in the Shuttle Flight Control Room (WFCR) in Johnson Space Center’s (JSC) Mission Control Center (MCC) with the STS-114 crewmembers during a fully-integrated simulation on October 13. The seven member crew was in a JSC-based simulator during the sims. The dress rehearsal of Discovery's rendezvous and docking with the International Space Station (ISS) was the first flight-specific training for the Space Shuttle's return to flight.

Yurchikhin and Parmitano in U.S. Laboratory

NASA Image and Video Library

2013-09-18

ISS037-E-001901 (18 Sept. 2013) --- In the International Space Station’s Destiny laboratory, Russian cosmonaut Fyodor Yurchikhin (right), Expedition 37 commander; and European Space Agency astronaut Luca Parmitano, flight engineer, watch the launch of the Orbital Sciences Corporation Antares rocket, with the Cygnus cargo spacecraft aboard, from Pad-0A of the Mid-Atlantic Regional Spaceport (MARS) NASA Wallops Flight Facility, Virginia. Cygnus is on its way to rendezvous with the space station and will deliver about 1,300 pounds (589 kilograms) of cargo, including food and clothing, to the Expedition 37 crew.

Nuclear electric propulsion mission engineering study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1973-01-01

Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied. The NEP stage design provides both inherent reliability and high payload mass capability. The NEP stage and payload integration was found to be compatible with the space shuttle.

U.S. Commercial Cargo Craft Heads to the Space Station

NASA Image and Video Library

2018-05-21

The remotely piloted Orbital ATK Cygnus cargo spacecraft launched May 21 from NASA's Wallops Flight Facility, Virginia atop an Antares rocket, headed for a rendezvous with the International Space Station to deliver several tons of scientific experiments and supplies for the station residents. Dubbed the SS “J.R. Thompson” in honor of the late spacefaring manager for both NASA and Orbital ATK, Cygnus will be robotically captured and installed to the earth-facing port of the station’s Unity module for a two-month stay at the orbital outpost.

PRELAUNCH ACTIVITY (GT-6) - ASTRONAUT THOMAS P. STAFFORD - MISC.

NASA Image and Video Library

1965-12-15

S65-61806 (15 Dec. 1965) --- Astronaut Thomas P. Stafford, Gemini-6 prime crew pilot, is seen through spacecraft window as he awaits the remaining minutes of the Gemini-6 prelaunch countdown. A two-day mission in space was scheduled for astronauts Stafford and Walter M. Schirra Jr. (out of frame), command pilot. NASA successfully launched Gemini-6 from Pad 19 at 8:37 a.m. (EST) on Dec. 15, 1965. An attempt will be made to rendezvous Gemini-6 with Gemini-7. Photo credit: NASA or National Aeronautics and Space Administration

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089829 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089793 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089802 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089782 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089830 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089820 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

ATV Fly-Under

NASA Image and Video Library

2014-08-08

ISS040-E-089798 (8 Aug. 2014) --- The “Georges Lemaitre” Automated Transfer Vehicle (ATV-5), photographed by an Expedition 40 crew member, flies directly under the International Space Station at a distance of about 3.7 miles to test sensors and radar systems designed for future European spacecraft. After its “fly-under” of the station, the ATV will move in front of, above, and behind the outpost for the final days of its two-week rendezvous that will lead to an automated docking to the aft port of the Zvezda Service Module on Aug. 12.

Validation of GNSS Multipath Model for Space Proximity Operations Using the Hubble Servicing Mission 4 Experiment

NASA Technical Reports Server (NTRS)

Ashman, B. W.; Veldman, J. L.; Axelrad, P.; Garrison, J. L.; Winternitz, L. B.

2016-01-01

In the rendezvous and docking of spacecraft, GNSS signals can reflect off the target vehicle and cause large errors in the chaser vehicle receiver at ranges below a few hundred meters. It has been proposed that the additional ray paths, or multipath, be used as a source of information about the state of the target relative to the receiver. With Hubble Servicing Mission 4 as a case study, electromagnetic ray tracing has been used to construct a model of reflected signals from known geometry. Oscillations in the prompt correlator power due to multipath, known as multipath fading, are studied as a means of model validation. Agreement between the measured and simulated multipath fading serves to confirm the presence of signals reflected off the target spacecraft that might be used for relative navigation.

Validation of GNSS Multipath Model for Space Proximity Operations Using the Hubble Servicing Mission 4 Experiment

NASA Technical Reports Server (NTRS)

Ashman, Ben; Veldman, Jeanette; Axelrad, Penina; Garrison, James; Winternitz, Luke

2016-01-01

In the rendezvous and docking of spacecraft, GNSS signals can reflect off the target vehicle and cause prohibitively large errors in the chaser vehicle receiver at ranges below 200 meters. It has been proposed that the additional ray paths, or multipath, be used as a source of information about the state of the target relative to the receiver. With Hubble Servicing Mission 4 as a case study, electromagnetic ray tracing has been used to construct a model of reflected signals from known geometry. Oscillations in the prompt correlator power due to multipath, known as multipath fading, are studied as a means of model validation. Agreement between the measured and simulated multipath fading serves to confirm the presence of signals reflected off the target spacecraft that might be used for relative navigation.

KSC-05PD-0133

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. From the nearby Press Site at Cape Canaveral Air Force Station, Fla., photographers capture the exciting launch of the Deep Impact spacecraft at 1:47 p.m. EST. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0134

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Erupting from the flames and smoke beneath it, NASAs Deep Impact spacecraft lifts off at 1:47 p.m. EST today from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0131

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Erupting from the flames and smoke beneath it, NASAs Deep Impact spacecraft lifts off at 1:47 p.m. EST today from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0135

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Erupting from the flames and smoke beneath it, NASAs Deep Impact spacecraft lifts off at 1:47 p.m. EST today from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0136

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Engulfed by flames and smoke, NASAs Deep Impact spacecraft lifts off at 1:47 p.m. EST today from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0130

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. With a burst of flames, NASAs Deep Impact spacecraft lifts off at 1:47 p.m. EST today from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

Cost, capability, and risk for planetary operations

NASA Technical Reports Server (NTRS)

Mclaughlin, William I.; Deutsch, Marie J.; Miller, Lanny J.; Wolff, Donna M.; Zawacki, Steven J.

1992-01-01

The three key factors for flight projects - cost, capability, and risk - are examined with respect to their interplay, the uplink process, cost drivers, and risk factors. Scientific objectives are translated into a computer program during the uplink process, and examples are given relating to the Voyager Interstellar Mission, Galileo, and the Comet Rendezvous Asteroid Flyby. The development of a multimission sequence system based on these uplinks is described with reference to specific subsystems such as the pointer and the sequence generator. Operational cost drivers include mission, flight-system, and ground-system complexity, uplink traffic, and work force. Operational risks are listed in terms of the mission operations, the environment, and the mission facilities. The uplink process can be analyzed in terms of software development, and spacecraft operability is shown to be an important factor from the initial stages of spacecraft development.

Large scale nonlinear programming for the optimization of spacecraft trajectories

NASA Astrophysics Data System (ADS)

Arrieta-Camacho, Juan Jose

Despite the availability of high fidelity mathematical models, the computation of accurate optimal spacecraft trajectories has never been an easy task. While simplified models of spacecraft motion can provide useful estimates on energy requirements, sizing, and cost; the actual launch window and maneuver scheduling must rely on more accurate representations. We propose an alternative for the computation of optimal transfers that uses an accurate representation of the spacecraft dynamics. Like other methodologies for trajectory optimization, this alternative is able to consider all major disturbances. In contrast, it can handle explicitly equality and inequality constraints throughout the trajectory; it requires neither the derivation of costate equations nor the identification of the constrained arcs. The alternative consist of two steps: (1) discretizing the dynamic model using high-order collocation at Radau points, which displays numerical advantages, and (2) solution to the resulting Nonlinear Programming (NLP) problem using an interior point method, which does not suffer from the performance bottleneck associated with identifying the active set, as required by sequential quadratic programming methods; in this way the methodology exploits the availability of sound numerical methods, and next generation NLP solvers. In practice the methodology is versatile; it can be applied to a variety of aerospace problems like homing, guidance, and aircraft collision avoidance; the methodology is particularly well suited for low-thrust spacecraft trajectory optimization. Examples are presented which consider the optimization of a low-thrust orbit transfer subject to the main disturbances due to Earth's gravity field together with Lunar and Solar attraction. Other example considers the optimization of a multiple asteroid rendezvous problem. In both cases, the ability of our proposed methodology to consider non-standard objective functions and constraints is illustrated. Future research directions are identified, involving the automatic scheduling and optimization of trajectory correction maneuvers. The sensitivity information provided by the methodology is expected to be invaluable in such research pursuit. The collocation scheme and nonlinear programming algorithm presented in this work, complement other existing methodologies by providing reliable and efficient numerical methods able to handle large scale, nonlinear dynamic models.

Launch and Assembly Reliability Analysis for Mars Human Space Exploration Missions

NASA Technical Reports Server (NTRS)

Cates, Grant R.; Stromgren, Chel; Cirillo, William M.; Goodliff, Kandyce E.

2013-01-01

NASA s long-range goal is focused upon human exploration of Mars. Missions to Mars will require campaigns of multiple launches to assemble Mars Transfer Vehicles in Earth orbit. Launch campaigns are subject to delays, launch vehicles can fail to place their payloads into the required orbit, and spacecraft may fail during the assembly process or while loitering prior to the Trans-Mars Injection (TMI) burn. Additionally, missions to Mars have constrained departure windows lasting approximately sixty days that repeat approximately every two years. Ensuring high reliability of launching and assembling all required elements in time to support the TMI window will be a key enabler to mission success. This paper describes an integrated methodology for analyzing and improving the reliability of the launch and assembly campaign phase. A discrete event simulation involves several pertinent risk factors including, but not limited to: manufacturing completion; transportation; ground processing; launch countdown; ascent; rendezvous and docking, assembly, and orbital operations leading up to TMI. The model accommodates varying numbers of launches, including the potential for spare launches. Having a spare launch capability provides significant improvement to mission success.

Incorporating Uncertainty into Spacecraft Mission and Trajectory Design

NASA Astrophysics Data System (ADS)

Juliana D., Feldhacker

The complex nature of many astrodynamic systems often leads to high computational costs or degraded accuracy in the analysis and design of spacecraft missions, and the incorporation of uncertainty into the trajectory optimization process often becomes intractable. This research applies mathematical modeling techniques to reduce computational cost and improve tractability for design, optimization, uncertainty quantication (UQ) and sensitivity analysis (SA) in astrodynamic systems and develops a method for trajectory optimization under uncertainty (OUU). This thesis demonstrates the use of surrogate regression models and polynomial chaos expansions for the purpose of design and UQ in the complex three-body system. Results are presented for the application of the models to the design of mid-eld rendezvous maneuvers for spacecraft in three-body orbits. The models are shown to provide high accuracy with no a priori knowledge on the sample size required for convergence. Additionally, a method is developed for the direct incorporation of system uncertainties into the design process for the purpose of OUU and robust design; these methods are also applied to the rendezvous problem. It is shown that the models can be used for constrained optimization with orders of magnitude fewer samples than is required for a Monte Carlo approach to the same problem. Finally, this research considers an application for which regression models are not well-suited, namely UQ for the kinetic de ection of potentially hazardous asteroids under the assumptions of real asteroid shape models and uncertainties in the impact trajectory and the surface material properties of the asteroid, which produce a non-smooth system response. An alternate set of models is presented that enables analytic computation of the uncertainties in the imparted momentum from impact. Use of these models for a survey of asteroids allows conclusions to be drawn on the eects of an asteroid's shape on the ability to successfully divert the asteroid via kinetic impactor.

The primer vector in linear, relative-motion equations. [spacecraft trajectory optimization

NASA Technical Reports Server (NTRS)

1980-01-01

Primer vector theory is used in analyzing a set of linear, relative-motion equations - the Clohessy-Wiltshire equations - to determine the criteria and necessary conditions for an optimal, N-impulse trajectory. Since the state vector for these equations is defined in terms of a linear system of ordinary differential equations, all fundamental relations defining the solution of the state and costate equations, and the necessary conditions for optimality, can be expressed in terms of elementary functions. The analysis develops the analytical criteria for improving a solution by (1) moving any dependent or independent variable in the initial and/or final orbit, and (2) adding intermediate impulses. If these criteria are violated, the theory establishes a sufficient number of analytical equations. The subsequent satisfaction of these equations will result in the optimal position vectors and times of an N-impulse trajectory. The solution is examined for the specific boundary conditions of (1) fixed-end conditions, two-impulse, and time-open transfer; (2) an orbit-to-orbit transfer; and (3) a generalized rendezvous problem. A sequence of rendezvous problems is solved to illustrate the analysis and the computational procedure.

A Delta-V map of the known Main Belt Asteroids

NASA Astrophysics Data System (ADS)

Taylor, Anthony; McDowell, Jonathan C.; Elvis, Martin

2018-05-01

With the lowered costs of rocket technology and the commercialization of the space industry, asteroid mining is becoming both feasible and potentially profitable. Although the first targets for mining will be the most accessible near Earth objects (NEOs), the Main Belt contains 106 times more material by mass. The large scale expansion of this new asteroid mining industry is contingent on being able to rendezvous with Main Belt asteroids (MBAs), and so on the velocity change required of mining spacecraft (delta-v). This paper develops two different flight burn schemes, both starting from Low Earth Orbit (LEO) and ending with a successful MBA rendezvous. These methods are then applied to the ∼700,000 asteroids in the Minor Planet Center (MPC) database with well-determined orbits to find low delta-v mining targets among the MBAs. There are 3986 potential MBA targets with a delta-v < 8 km s-1 , but the distribution is steep and reduces to just 4 with delta-v < 7 km s-1. The two burn methods are compared and the orbital parameters of low delta-v MBAs are explored.

A CubeSat Asteroid Mission: Design Study and Trade-Offs

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Oleson, Steven R.; McGuire, Melissa; Hepp, Aloysius; Stegeman, James; Bur, Mike; Burke, Laura; Martini, Michael; Fittje, James E.; Kohout, Lisa;

On the Shoulders of Titans: A History of Project Gemini

NASA Technical Reports Server (NTRS)

Hacker, B. C.

1977-01-01

Gemini was the intermediate manned space flight program between America's first steps into space with Mercury and the manned lunar expeditions of Apollo. Because of its position between these two other efforts, Gemini is probably less remembered. Still, it more than had its place in man's progress into this new frontier. Gemini accomplishments were manyfold. They included many firsts: first astronaut-controlled maneuvering in space; first rendezvous in space of one spacecraft with another; first docking of one spacecraft with a propulsive stage and use of that stage to transfer man to high altitude; first traverse of man into the earth's radiation belts; first extended manned flights of a week or more in duration; first extended stays of man outside his spacecraft; first controlled reentry and precision landing; and many more. These achievements were significant in ways one cannot truly evaluate even today, but two things stand out: (1) it was the time when America caught up and surpassed the Soviet Union in manned space flight, and (2) these demonstrations of capability were an absolute prerequisite to the phenomenal Apollo accomplishments then yet to come.

Capture of near-Earth objects with low-thrust propulsion and invariant manifolds

NASA Astrophysics Data System (ADS)

Tang, Gao; Jiang, Fanghua

2016-01-01

In this paper, a mission incorporating low-thrust propulsion and invariant manifolds to capture near-Earth objects (NEOs) is investigated. The initial condition has the spacecraft rendezvousing with the NEO. The mission terminates once it is inserted into a libration point orbit (LPO). The spacecraft takes advantage of stable invariant manifolds for low-energy ballistic capture. Low-thrust propulsion is employed to retrieve the joint spacecraft-asteroid system. Global optimization methods are proposed for the preliminary design. Local direct and indirect methods are applied to optimize the two-impulse transfers. Indirect methods are implemented to optimize the low-thrust trajectory and estimate the largest retrievable mass. To overcome the difficulty that arises from bang-bang control, a homotopic approach is applied to find an approximate solution. By detecting the switching moments of the bang-bang control the efficiency and accuracy of numerical integration are guaranteed. By using the homotopic approach as the initial guess the shooting function is easy to solve. The relationship between the maximum thrust and the retrieval mass is investigated. We find that both numerically and theoretically a larger thrust is preferred.

Deep Space 1 Using its Ion Engine Artist Concept

NASA Image and Video Library

2003-07-02

NASA's New Millennium Deep Space 1 spacecraft approaching the comet 19P/Borrelly. With its primary mission to serve as a technology demonstrator--testing ion propulsion and 11 other advanced technologies--successfully completed in September 1999, Deep Space 1 is now headed for a risky, exciting rendezvous with Comet Borrelly. NASA extended the mission, taking advantage of the ion propulsion and other systems to target the daring encounter with the comet in September 2001. Once a sci-fi dream, the ion propulsion engine has powered the spacecraft for over 12,000 hours. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The first flight in NASA's New Millennium Program, Deep Space 1 was launched October 24, 1998 aboard a Boeing Delta 7326 rocket from Cape Canaveral Air Station, FL. Deep Space 1 successfully completed and exceeded its mission objectives in July 1999 and flew by a near-Earth asteroid, Braille (1992 KD), in September 1999. http://photojournal.jpl.nasa.gov/catalog/PIA04604

Mission Architecture Comparison for Human Lunar Exploration

NASA Technical Reports Server (NTRS)

Geffre, Jim; Robertson, Ed; Lenius, Jon

2006-01-01

The Vision for Space Exploration outlines a bold new national space exploration policy that holds as one of its primary objectives the extension of human presence outward into the Solar System, starting with a return to the Moon in preparation for the future exploration of Mars and beyond. The National Aeronautics and Space Administration is currently engaged in several preliminary analysis efforts in order to develop the requirements necessary for implementing this objective in a manner that is both sustainable and affordable. Such analyses investigate various operational concepts, or mission architectures , by which humans can best travel to the lunar surface, live and work there for increasing lengths of time, and then return to Earth. This paper reports on a trade study conducted in support of NASA s Exploration Systems Mission Directorate investigating the relative merits of three alternative lunar mission architecture strategies. The three architectures use for reference a lunar exploration campaign consisting of multiple 90-day expeditions to the Moon s polar regions, a strategy which was selected for its high perceived scientific and operational value. The first architecture discussed incorporates the lunar orbit rendezvous approach employed by the Apollo lunar exploration program. This concept has been adapted from Apollo to meet the particular demands of a long-stay polar exploration campaign while assuring the safe return of crew to Earth. Lunar orbit rendezvous is also used as the baseline against which the other alternate concepts are measured. The first such alternative, libration point rendezvous, utilizes the unique characteristics of the cislunar libration point instead of a low altitude lunar parking orbit as a rendezvous and staging node. Finally, a mission strategy which does not incorporate rendezvous after the crew ascends from the Moon is also studied. In this mission strategy, the crew returns directly to Earth from the lunar surface, and is thus referred to as direct return. Figures of merit in the areas of safety and mission success, mission effectiveness, extensibility, and affordability are used to evaluate and compare the lunar orbit rendezvous, libration point rendezvous, and direct return architectures, and this paper summarizes the results of those assessments.

ASTP (SA-210) Launch vehicle operational flight trajectory. Part 3: Final documentation

NASA Technical Reports Server (NTRS)

Carter, A. B.; Klug, G. W.; Williams, N. W.

1975-01-01

Trajectory data are presented for a nominal and two launch window trajectory simulations. These trajectories are designed to insert a manned Apollo spacecraft into a 150/167 km. (81/90 n. mi.) earth orbit inclined at 51.78 degrees for rendezvous with a Soyuz spacecraft, which will be orbiting at approximately 225 km. (121.5 n. mi.). The launch window allocation defined for this launch is 500 pounds of S-IVB stage propellant. The launch window opening trajectory simulation depicts the earliest launch time deviation from a planar flight launch which conforms to this constraint. The launch window closing trajectory simulation was developed for the more stringent Air Force Eastern Test Range (AFETR) flight azimuth restriction of 37.4 degrees east-of-north. These trajectories enclose a 12.09 minute launch window, pertinent features of which are provided in a tabulation. Planar flight data are included for mid-window reference.

KSC-108-75P-0057

NASA Image and Video Library

1975-02-10

CAPE CANAVERAL, Fla. – The Soviet and American crews for the July Apollo Soyuz Test Project [standing, center] addressed personnel assembled in a firing room at KSC on February 10. The crews for the joint manned space mission toured the Center during their three-day visit which also included inspection of ASTP equipment and facilities and a trip to Disney World. The first international crewed spaceflight was a joint U.S.-U.S.S.R. rendezvous and docking mission. The Apollo-Soyuz Test Project, or ASTP, took its name from the spacecraft employed: the American Apollo and the Soviet Soyuz. The three-man Apollo crew lifted off from Kennedy Space Center aboard a Saturn IB rocket on July 15, 1975, to link up with the Soyuz that had launched a few hours earlier. A cylindrical docking module served as an airlock between the two spacecraft for transfer of the crew members. Photo credit: NASA

Pointing control for the International Comet Mission

NASA Technical Reports Server (NTRS)

Leblanc, D. R.; Schumacher, L. L.

1980-01-01

The design of the pointing control system for the proposed International Comet Mission, intended to fly by Comet Halley and rendezvous with Comet Tempel-2 is presented. Following a review of mission objectives and the spacecraft configuration, design constraints on the pointing control system controlling the two-axis gimballed scan platform supporting the science instruments are discussed in relation to the scientific requirements of the mission. The primary design options considered for the pointing control system design for the baseline spacecraft are summarized, and the design selected, which employs a target-referenced, inertially stabilized control system, is described in detail. The four basic modes of operation of the pointing control subsystem (target acquisition, inertial hold, target track and slew) are discussed as they relate to operations at Halley and Tempel-2. It is pointed that the pointing control system design represents a significant advance in the state of the art of pointing controls for planetary missions.

KSC-05PP-0138

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Emerging through the smoke and steam, the Boeing Delta II rocket carrying NASAs Deep Impact spacecraft lifts off at 1:47 p.m. EST from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0137

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. After a perfect liftoff at 1:47 p.m. EST today from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., the Boeing Delta II rocket with Deep Impact spacecraft aboard soars through the clear blue sky. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0128

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. On Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., the Boeing Delta II rocket carrying the Deep Impact spacecraft stands out against an early dawn sky. Scheduled for liftoff at 1:47 p.m. EST today, Deep Impact will head for space and a rendezvous with Comet Tempel 1 when the comet is 83 million miles from Earth. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network. Deep Impact is a NASA Discovery mission.

KSC-05pp0138

NASA Image and Video Library

2005-01-12

KENNEDY SPACE CENTER, FLA. - Emerging through the smoke and steam, the Boeing Delta II rocket carrying NASA’s Deep Impact spacecraft lifts off at 1:47 p.m. EST from Launch Pad 17-B, Cape Canaveral Air Force Station, Fla. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impact’s flyby spacecraft will reveal the secrets of the comet’s interior by collecting pictures and data of how the crater forms, measuring the crater’s depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

KSC-05PD-0124

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. On Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., the Boeing Delta II rocket carrying the Deep Impact spacecraft is bathed in light waiting for tower rollback before launch. Scheduled for liftoff at 1:47 p.m. EST today, Deep Impact will head for space and a rendezvous with Comet Tempel 1 when the comet is 83 million miles from Earth. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network. Deep Impact is a NASA Discovery mission.

KSC-05PD-0132

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. Guests of NASA gather near the launch site at Cape Canaveral Air Force Station, Fla., to watch the Deep Impact spacecraft as it speeds through the air after a perfect launch at 1:47 p.m. EST. A NASA Discovery mission, Deep Impact is heading for space and a rendezvous 83 million miles from Earth with Comet Tempel 1. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network.

Target selection for a hypervelocity asteroid intercept vehicle flight validation mission

NASA Astrophysics Data System (ADS)

Wagner, Sam; Wie, Bong; Barbee, Brent W.

2015-02-01

Asteroids and comets have collided with the Earth in the past and will do so again in the future. Throughout Earth's history these collisions have played a significant role in shaping Earth's biological and geological histories. The planetary defense community has been examining a variety of options for mitigating the impact threat of asteroids and comets that approach or cross Earth's orbit, known as near-Earth objects (NEOs). This paper discusses the preliminary study results of selecting small (100-m class) NEO targets and mission analysis and design trade-offs for validating the effectiveness of a Hypervelocity Asteroid Intercept Vehicle (HAIV) concept, currently being investigated for a NIAC (NASA Advanced Innovative Concepts) Phase 2 study. In particular this paper will focus on the mission analysis and design for single spacecraft direct impact trajectories, as well as several mission types that enable a secondary rendezvous spacecraft to observe the HAIV impact and evaluate it's effectiveness.

Design concepts and performance of NASA X-band transponder (DST) for deep space spacecraft applications

NASA Technical Reports Server (NTRS)

Mysoor, Narayan R.; Perret, Jonathan D.; Kermode, Arthur W.

1991-01-01

The design concepts and measured performance characteristics of an X band (7162 MHz/8415 MHz) breadboard deep space transponder (DST) for future spacecraft applications, with the first use scheduled for the Comet Rendezvous Asteroid Flyby (CRAF) and Cassini missions in 1995 and 1996, respectively. The DST consists of a double conversion, superheterodyne, automatic phase tracking receiver, and an X band (8415 MHz) exciter to drive redundant downlink power amplifiers. The receiver acquires and coherently phase tracks the modulated or unmodulated X band (7162 MHz) uplink carrier signal. The exciter phase modulates the X band (8415 MHz) downlink signal with composite telemetry and ranging signals. The receiver measured tracking threshold, automatic gain control static phase error, and phase jitter characteristics of the breadboard DST are in good agreement with the expected performance. The measured results show a receiver tracking threshold of -158 dBm and a dynamic signal range of 88 dB.

An X-band spacecraft transponder for deep space applications - Design concepts and breadboard performance

NASA Technical Reports Server (NTRS)

Mysoor, Narayan R.; Perret, Jonathan D.; Kermode, Arthur W.

1992-01-01

The design concepts and measured performance characteristics are summarized of an X band (7162 MHz/8415 MHz) breadboard deep space transponder (DSP) for future spacecraft applications, with the first use scheduled for the Comet Rendezvous Asteroid Flyby (CRAF) and Cassini missions in 1995 and 1996, respectively. The DST consists of a double conversion, superheterodyne, automatic phase tracking receiver, and an X band (8415 MHz) exciter to drive redundant downlink power amplifiers. The receiver acquires and coherently phase tracks the modulated or unmodulated X band (7162 MHz) uplink carrier signal. The exciter phase modulates the band (8415 MHz) downlink signal with composite telemetry and ranging signals. The receiver measured tracking threshold, automatic gain control, static phase error, and phase jitter characteristics of the breadboard DST are in good agreement with the expected performance. The measured results show a receiver tracking threshold of -158 dBm and a dynamic signal range of 88 dB.

Optimal trajectories for the aeroassisted flight experiment. Part 3: Formulation, results, and analysis

NASA Technical Reports Server (NTRS)

Miele, A.; Wang, T.; Lee, W. Y.; Zhao, Z. G.

1989-01-01

The determination of optimal trajectories for the aero-assisted flight experiment (AFE) is investigated. The intent of this experiment is to simulate a GEO-to-LEO transfer, where GEO denotes a geosynchronous Earth orbit and LEO denotes a low Earth orbit. The trajectories of an AFE spacecraft are analyzed in a 3D-space, employing the full system of 6 ODEs describing the atmospheric pass. The atmospheric entry conditions are given, and the atmospheric exit conditions are adjusted in such a way that the following conditions are satisfied: (1) the atmospheric velocity depletion is such that, after exiting, the AFE spacecraft first ascends to a specified apogee and then descends to a specified perigee; and (2) the exit orbital plane is identical with the entry orbital plane. The final maneuver, not analyzed here, includes the rendezvous with and the capture by the space shuttle.

Kuiper Belt Object Orbiter Using Advanced Radioisotope Power Sources and Electric Propulsion

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; McGuire, Melissa L.; Dankanich, John; Colozza, Anthony; Schmitz, Paul; Khan, Omair; Drexler, Jon; Fittje, James

2011-01-01

A joint NASA GRC/JPL design study was performed for the NASA Radioisotope Power Systems Office to explore the use of radioisotope electric propulsion for flagship class missions. The Kuiper Belt Object Orbiter is a flagship class mission concept projected for launch in the 2030 timeframe. Due to the large size of a flagship class science mission larger radioisotope power system building blocks were conceptualized to provide the roughly 4 kW of power needed by the NEXT ion propulsion system and the spacecraft. Using REP the spacecraft is able to rendezvous with and orbit a Kuiper Belt object in 16 years using either eleven (no spare) 420 W advanced RTGs or nine (with a spare) 550 W advanced Stirling Radioisotope systems. The design study evaluated integrating either system and estimated impacts on cost as well as required General Purpose Heat Source requirements.

Earth Science

NASA Image and Video Library

1996-01-13

The Near Earth Asteroid Rendezvous (NEAR) spacecraft undergoing preflight preparation in the Spacecraft Assembly Encapsulation Facility-2 (SAEF-2) at Kennedy Space Center (KSC). NEAR will perform two critical mission events - Mathilde flyby and the Deep-Space maneuver. NEAR will fly-by Mathilde, a 38-mile (61-km) diameter C-type asteroid, making use of its imaging system to obtain useful optical navigation images. The primary science instrument will be the camera, but measurements of magnetic fields and mass also will be made. The Deep-Space Maneuver (DSM) will be executed about a week after the Mathilde fly-by. The DSM represents the first of two major burns during the NEAR mission of the 100-pound bi-propellant (Hydrazine/nitrogen tetroxide) thruster. This maneuver is necessary to lower the perihelion distance of NEAR's trajectory. The DSM will be conducted in two segments to minimize the possibility of an overburn situation.

President Ford and both the Soviet and American ASTP crews

NASA Technical Reports Server (NTRS)

1974-01-01

President Gerald R. Ford removes the Soviet Soyuz spacecraft model from a model set depicting the 1975 Apollo Soyuz Test Project (ASTP), an Earth orbital docking and rendezvous mission with crewmen from the U.S. and USSR. From left to right, Vladamir A. Shatalov, Chief, Cosmonaut training; Valeriy N. Kubasov, ASTP Soviet engineer; Aleksey A. Leonov, ASTP Soviet crew commander; Thomas P. Stafford, commander of the American crew; Donald K. Slayton, American docking module pilot; Vance D. Brand, command module pilot for the American crew. Dr. George M Low, Deputy Administrator for NASA is partially obscured behind President Ford.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This illustration is the Lunar Module (LM) configuration. The LM was a two part spacecraft. Its lower or descent stage had the landing gear, engines, and fuel needed for the landing. When the LM blasted off the Moon, the descent stage served as the launching pad for its companion ascent stage, which was also home for the two astronauts on the surface of the Moon. The LM was full of gear with which to communicate, navigate, and rendezvous. It also had its own propulsion system, and an engine to lift it off the Moon and send it on a course toward the orbiting Command Module.

KSC-99pc49

NASA Image and Video Library

1999-01-11

In the Payload Hazardous Servicing Facility, workers look over the solar panels on the Stardust spacecraft that are deployed for lighting tests. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule to be jettisoned as it swings by Earth in January 2006

KSC-99pc38

NASA Image and Video Library

1999-01-11

Workers in the Payload Hazardous Servicing Facility deploy a solar panel on the Stardust spacecraft before performing lighting tests. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule to be jettisoned as it swings by Earth in January 2006

ART CONCEPTS - APOLLO-SOYUZ TEST PROJECT (ASTP)

NASA Image and Video Library

1975-04-01

S75-27288 (April 1975) --- An artist?s concept illustrating the mission profile of the Apollo-Soyuz Test Project. The phases of the mission depicted include launch, rendezvous, docking, separation and splashdown. During the joint U.S.-USSR ASTP flight, scheduled for July 1975, the American and Soviet crews will visit one another?s spacecraft while the Soyuz and Apollo are docked for a maximum period of two days. The mission is designed to test equipment and techniques that will establish international crew rescue capability in space, as well as permit future cooperative scientific missions. This artwork is by Davis Meltzer.

Saturn/Titan Rendezvous: A Solar-Sail Aerocapture Mission

NASA Technical Reports Server (NTRS)

Matloff, Gregory L.; Taylor, Travis; Powell, Conley

2004-01-01

A low-mass Titan orbiter is proposed that uses conservative or optimistic solar sails for all post-Earth-escape propulsion. After accelerating the probe onto a trans-Saturn trajectory, the sail is used parachute style for Saturn capture during a pass through Saturn's outer atmosphere. If the apoapsis of the Saturn-capture orbit is appropriate, the aerocapture maneuver can later be repeated at Titan so that the spacecraft becomes a satellite of Titan. An isodensity-atmosphere model is applied to screen aerocapture trajectories. Huygens/Cassini should greatly reduce uncertainties regarding the upper atmospheres of Saturn and Titan.

Origins, Spectral Interpretation, Resource Identification, Security, Regolith Explorer Planning (OSIRIS-REx)

NASA Technical Reports Server (NTRS)

Nakamura-Messenger, Keiko; Messenger, Scott; Keller, Lindsay; Righter, Kevin

2014-01-01

Scientists at ARES are preparing to curate and analyze samples from the first U.S. mission to return samples from an asteroid. The Origins-Spectral Interpretation- Resource Identification-Security-Regolith Explorer, or OSIRIS-REx, was selected by NASA as the third mission in its New Frontiers Program. The robotic spacecraft will launch in 2016 and rendezvous with the near-Earth asteroid Bennu, in 2020. A robotic arm will collect at least 60 grams of material from the surface of the asteroid to be returned to Earth in 2023 for worldwide distribution by the NASA Astromaterials Curation Facility at ARES.

Kotov practices the manual docking techniques with the TORU

NASA Image and Video Library

2013-11-22

ISS038-E-006656 (22 Nov. 2013) --- Russian cosmonaut Oleg Kotov, Expedition 38 commander, practices manual docking techniques with the TORU, or telerobotically operated rendezvous system, in the Zvezda Service Module of the International Space Station in preparation for the docking of the Progress 53 spacecraft. Kotov, using the Simvol-TS screen and hand controllers, could manually dock the Progress to the station in the event of a failure of the Kurs automated docking system. The Progress 53 craft is scheduled to complete its automated docking to the aft port of Zvezda at 5:28 p.m. (EST) on Nov. 29.

TORU OBT

NASA Image and Video Library

2014-07-22

ISS040-E-070857 (22 July 2014) --- Russian cosmonaut Alexander Skvortsov, Expedition 40 flight engineer, practices manual docking techniques with the TORU, or telerobotically operated rendezvous system, in the Zvezda Service Module of the International Space Station in preparation for the docking of the Progress 56 spacecraft. Skvortsov, using the Simvol-TS screen and hand controllers, could manually dock the Progress to the station in the event of a failure of the Kurs automated docking system. The Progress 56 craft is scheduled to complete its automated docking to the Pirs docking compartment at 11:30 p.m. (EDT) on July 23, 2014.

Tyurin practices the manual docking techniques with the TORU

NASA Image and Video Library

2013-11-22

ISS038-E-006663 (22 Nov. 2013) --- Russian cosmonaut Mikhail Tyurin, Expedition 38 flight engineer, practices manual docking techniques with the TORU, or telerobotically operated rendezvous system, in the Zvezda Service Module of the International Space Station in preparation for the docking of the Progress 53 spacecraft. Tyurin, using the Simvol-TS screen and hand controllers, could manually dock the Progress to the station in the event of a failure of the Kurs automated docking system. The Progress 53 craft is scheduled to complete its automated docking to the aft port of Zvezda at 5:28 p.m. (EST) on Nov. 29.

TORU OBT

NASA Image and Video Library

2014-07-22

ISS040-E-070859 (22 July 2014) --- Russian cosmonaut Alexander Skvortsov, Expedition 40 flight engineer, practices manual docking techniques with the TORU, or telerobotically operated rendezvous system, in the Zvezda Service Module of the International Space Station in preparation for the docking of the Progress 56 spacecraft. Skvortsov, using the Simvol-TS screen and hand controllers, could manually dock the Progress to the station in the event of a failure of the Kurs automated docking system. The Progress 56 craft is scheduled to complete its automated docking to the Pirs docking compartment at 11:30 p.m. (EDT) on July 23, 2014.

Automated and Adaptive Mission Planning for Orbital Express

NASA Technical Reports Server (NTRS)

Chouinard, Caroline; Knight, Russell; Jones, Grailing; Tran, Daniel; Koblick, Darin

2008-01-01

The Orbital Express space mission was a Defense Advanced Research Projects Agency (DARPA) lead demonstration of on-orbit satellite servicing scenarios, autonomous rendezvous, fluid transfers of hydrazine propellant, and robotic arm transfers of Orbital Replacement Unit (ORU) components. Boeing's Autonomous Space Transport Robotic Operations (ASTRO) vehicle provided the servicing to the Ball Aerospace's Next Generation Serviceable Satellite (NextSat) client. For communication opportunities, operations used the high-bandwidth ground-based Air Force Satellite Control Network (AFSCN) along with the relatively low-bandwidth GEO-Synchronous space-borne Tracking and Data Relay Satellite System (TDRSS) network. Mission operations were conducted out of the RDT&E Support Complex (RSC) at the Kirtland Air Force Base in New Mexico. All mission objectives were met successfully: The first of several autonomous rendezvous was demonstrated on May 5, 2007; autonomous free-flyer capture was demonstrated on June 22, 2007; the fluid and ORU transfers throughout the mission were successful. Planning operations for the mission were conducted by a team of personnel including Flight Directors, who were responsible for verifying the steps and contacts within the procedures, the Rendezvous Planners who would compute the locations and visibilities of the spacecraft, the Scenario Resource Planners (SRPs), who were concerned with assignment of communications windows, monitoring of resources, and sending commands to the ASTRO spacecraft, and the Mission planners who would interface with the real-time operations environment, process planning products and coordinate activities with the SRP. The SRP position was staffed by JPL personnel who used the Automated Scheduling and Planning ENvironment (ASPEN) to model and enforce mission and satellite constraints. The lifecycle of a plan began three weeks outside its execution on-board. During the planning timeframe, many aspects could change the plan, causing the need for re-planning. These variable factors, ranging from shifting contact times to ground-station closures and required maintenance times, are discussed along with the flexibility of the ASPEN tool to accommodate changes to procedures and the daily or long-range plan, which contributed to the success of the mission. This paper will present an introduction to ASPEN, a more in-depth discussion on its use on the Orbital Express mission, and other relative work. A description of ground operations after the SRP deliveries were made is included, and we briefly discuss lessons learned from the planning perspective and future work.

Determination of motion extrema in multi-satellite systems

NASA Astrophysics Data System (ADS)

Allgeier, Shawn E.

Spacecraft, or satellite formation flight has been a topic of interest dating back to the Gemini program of the 1960s. Traditionally space missions have been designed around large monolithic assets. Recent interest in low cost, rapid call up mission architectures structured around fractionated systems, small satellites, and constellations has spurred renewed efforts in spacecraft relative motion problems. While such fractionated, or multi-body systems may provide benefits in terms of risk mitigation and cost savings, they introduce new technical challenges in terms of satellite coordination. Characterization of satellite formations is a vital requirement for them to have utility to industry and government entities. Satellite formations introduce challenges in the form of constellation maintenance, inter-satellite communications, and the demand for more sophisticated guidance, navigation, and control systems. At the core of these challenges is the orbital mechanics which govern the resulting motion. New applications of algebraic techniques are applied to the formation flight problem, specifically Gröbner basis tools, as a means of determining extrema of certain quantities pertaining to formation flight. Specifically, bounds are calculated for the relative position components, relative speed, relative velocity components, and range rate. The position based metrics are relevant for planning formation geometry, particularly in constellation or Earth observation applications. The velocity metrics are relevant in the design of end game interactions for rendezvous and proximity operations. The range rate of one satellite to another is essential in the design of radio frequency hardware for inter-satellite communications so that the doppler shift can be calculated a priori. Range rate may also have utility in space based surveillance and space situational awareness concerns, such as cross tagging. The results presented constitute a geometric perspective and have utility to mission designers, particularly for missions involving rendezvous and proximity operations.

Science Planning for the TROPIX Mission

NASA Technical Reports Server (NTRS)

Russell, C. T.

1998-01-01

The objective of the study grant was to undertake the planning needed to execute meaningful solar electric propulsion missions in the magnetosphere and beyond. The first mission examined was the Transfer Orbit Plasma Investigation Experiment (TROPIX) mission to spiral outward through the magnetosphere. The next mission examined was to the moon and an asteroid. Entitled Diana, it was proposed to NASA in October 1994. Two similar missions were conceived in 1996 entitled CNR for Comet Nucleus Rendezvous and MBAR for Main Belt Asteroid Rendezvous. The latter mission was again proposed in 1998. All four of these missions were unsuccessfully proposed to the NASA Discovery program. Nevertheless we were partially successful in that the Deep Space 1 (DS1) mission was eventually carried out nearly duplicating our CNR mission. Returning to the magnetosphere we studied and proposed to the Medium Class Explorer (MIDEX) program a MidEx mission called TEMPEST, in 1995. This mission included two solar electric spacecraft that spiraled outward in the magnetosphere: one at near 900 inclination and one in the equatorial plane. This mission was not selected for flight. Next we proposed a single SEP vehicle to carry Energetic Neutral Atom (ENA) imagers and inside observations to complement the IMAGE mission providing needed data to properly interpret the IMAGE data. This mission called SESAME was submitted unsuccessfully in 1997. One proposal was successful. A study grant was awarded to examine a four spacecraft solar electric mission, named Global Magnetospheric Dynamics. This study was completed and a report on this mission is attached but events overtook this design and a separate study team was selected to design a classical chemical mission as a Solar Terrestrial Probe. Competing proposals such as through the MIDEX opportunity were expressly forbidden. A bibliography is attached.

Coordinated Radio, Electron, and Waves Experiment (CREWE) for the NASA Comet Rendezvous and Asteroid Flyby (CRAF) instrument

NASA Technical Reports Server (NTRS)

Scudder, Jack D.

1992-01-01

The Coordinated Radio, Electron, and Waves Experiment (CREWE) was designed to determine density, bulk velocity and temperature of the electrons for the NASA Comet Rendezvous and Asteroid Flyby Spacecraft, to define the MHD-SW IMF flow configuration; to clarify the role of impact ionization processes, to comment on the importance of anomalous ionization phenomena (via wave particle processes), to quantify the importance of wave turbulence in the cometary interaction, to establish the importance of photoionization via the presence of characteristic lines in a structured energy spectrum, to infer the presence and grain size of significant ambient dust column density, to search for the theoretically suggested 'impenetrable' contact surface, and to quantify the flow of heat (in the likelihood that no surface exists) that will penetrate very deep into the atmosphere supplying a good deal of heat via impact and charge exchange ionization. This final report provides an instrument description, instrument test plans, list of deliverables/schedule, flight and support equipment and software schedule, CREWE accommodation issues, resource requirements, status of major contracts, an explanation of the non-NASA funded efforts, status of EIP and IM plan, descope options, and Brinton questions.

Fourth Report of the Task Force on the Shuttle-Mir Rendezvous and Docking Missions

NASA Technical Reports Server (NTRS)

1995-01-01

On December 6, 1994, the NASA Administrator, Mr. Daniel Goldin, requested that Lt. Gen. Thomas P. Stafford, in his role as the Chairman of the NASA Advisory Council Task Force on the Shuttle-Mir Rendezvous and Docking Missions, lead a team composed of several Task Force members and technical advisors' to Russia with the goal of reviewing preparations and readiness for the upcoming international Space Station Phase 1 missions. In his directions to Gen. Stafford, Mr. Goldin requested that the review team focus its initial efforts on safety of flight issues for the following Phase 1A missions: the Soyuz TM-21 mission which will carry U.S. astronaut Dr. Norman Thagard and cosmonauts Lt. Col. Vladimir Dezhurov and Mr. Gennady Strekalov aboard a Soyuz spacecraft to the Mir Station; the Mir 18 Main Expedition during which Thagard and his fellow cosmonauts, Dezhurov and Strokalov, will spend approximately three months aboard the Mir Station; the STS-71 Space Shuttle mission which will perform the first Shuttle-Mir docking, carry cosmonauts Col. Anatoly SoloViev and Mr. Nikolai Budarin to the Mir Station, and return Thagard, Dezhurov, and Strekalov to Earth.

Can 67P/Churyumov-Gerasimenko become the reference for comet research?

NASA Astrophysics Data System (ADS)

Schulz, R.

2014-07-01

After its discovery in 1969, comet 67P/Churyumov-Gerasimenko went almost unnoticed through another five perihelion passages until the year 2003, when it suddenly became the new target of the first comet rendezvous mission, Rosetta. Today, 11 years and 1.5 apparitions later, it has become one of the few Jupiter-family comets that were monitored along its entire orbit, even near aphelion. Huge effort was spent in determining its characteristics. Observations obtained by the largest and most sophisticated telescopes on the Earth and in space were combined with dedicated modelling approaches in order to be best prepared for the space mission. Therefore, at this point time, we have basically determined as much as is achievable for a comet of this brightness without visiting it by spacecraft. A summary of what we already know about 67P/Churyumov-Gerasimenko, hence what we could in principle also determine for the ensemble of Jupiter-family comets, will be provided. The information expected to become available after the comet rendezvous will then be discussed particularly in view of whether and how it can be transferred to other comets for which only remote observations can be collected in the near future.

Fuzzy logic techniques for rendezvous and docking of two geostationary satellites

NASA Technical Reports Server (NTRS)

Ortega, Guillermo

1995-01-01

Large assemblings in space require the ability to manage rendezvous and docking operations. In future these techniques will be required for the gradual build up of big telecommunication platforms in the geostationary orbit. The paper discusses the use of fuzzy logic to model and implement a control system for the docking/berthing of two satellites in geostationary orbit. The system mounted in a chaser vehicle determines the actual state of both satellites and generates torques to execute maneuvers to establish the structural latching. The paper describes the proximity operations to collocate the two satellites in the same orbital window, the fuzzy guidance and navigation of the chaser approaching the target and the final Fuzzy berthing. The fuzzy logic system represents a knowledge based controller that realizes the close loop operations autonomously replacing the conventional control algorithms. The goal is to produce smooth control actions in the proximity of the target and during the docking to avoid disturbance torques in the final assembly orbit. The knowledge of the fuzzy controller consists of a data base of rules and the definitions of the fuzzy sets. The knowledge of an experienced spacecraft controller is captured into a set of rules forming the Rules Data Base.

Intrepid: Exploring the NEA population with a Fleet of Highly Autonomous SmallSat explorers

NASA Astrophysics Data System (ADS)

Blacksberg, Jordana; Chesley, Steven R.; Ehlmann, Bethany; Raymond, Carol Anne

2017-10-01

The Intrepid mission concept calls for phased deployment of a fleet of small highly autonomous rendezvous spacecraft designed to characterize the evolution, structure and composition of dozens of Near-Earth Asteroids (NEAs). Intrepid represents a marked departure from conventional solar system exploration projects, where a single unique and complex spacecraft is typically directed to explore a single target body. In contrast, Intrepid relies on the deployment of a large number of autonomous spacecraft to provide redundancy and ensure that the project goals are achieved at a small fraction of the cost of typical missions.The Intrepid science goals are threefold: (1) to understand the evolutionary processes that govern asteroid physical, chemical and dynamical histories and relate these results to solar system origins and evolution; (2) to facilitate impactor deflection scenarios for planetary defense by statistically characterizing relevant asteroid physical properties; (3) to quantify the presence and extractability of potentially useful resources on a large sample of asteroids. To achieve these goals, the baseline architecture includes multiple modular instruments including cameras, spectrometers, radar sounders, and projectiles that could interact with the target asteroid. Key questions to be addressed are: what is the total quantity of water in each object? How is the water incorporated? Are organics present? What is the asteroid physical structure? How would the object respond to impact/deflection?We have begun development of a miniature infrared point spectrometer, a cornerstone of the Intrepid payload, covering both shortwave infrared (SWIR) and mid-infrared (MIR) spectral bands. The spectrometer is designed with a compact 2U form-factor, making it both relevant to Intrepid and implementable on a CubeSat. The combination of SWIR and MIR in a single integrated instrument would enable robust compositional interpretations from a single dataset combining both solar reflectance and thermal emission spectroscopy. These measurements would be crucial to determining the quantity and nature of water present.

Apollo-Lunar Orbital Rendezvous Technique

NASA Technical Reports Server (NTRS)

1963-01-01

The film shows artists rendition of the spacecrafts, boosters, and flight of the Apollo lunar missions. The Apollo spacecraft will consist of three modules: the manned Command Module; the Service Module, which contains propulsion systems; and the Lunar Excursion Module (LEM) to carry astronauts to the moon and back to the Command and Service Modules. The spacecraft will be launched via a three-stage Saturn booster. The first stage will provide 7.5 million pounds of thrust from five F-1 engines for liftoff and initial powered flight. The second stage will develop 1 million pounds of thrust from five J-2 engines to boost the spacecraft almost into Earth orbit. Immediately after ignition of the second stage, the Launch Escape System will be jettisoned. A single J-2 engine in the S4B stage will provide 200,000 pounds of thrust to place the spacecraft in an earth parking orbit. It also will be used to propel the spacecraft into a translunar trajectory, then it will separate from the Apollo Modules. Onboard propulsion systems will be used to insert the spacecraft into lunar orbit. Two astronauts will enter the LEM, which will separate from the command and service modules. The LEM will go into elliptical orbit and prepare for landing. The LEM will lift off of the Moon's surface to return to the Command and Service Modules, and most likely be left in lunar orbit. After leaving the Moon's orbit, and shortly before entering Earth's orbit, the Service Module will be ejected. The Command Module will be oriented for reentry into the Earth's atmosphere. A drogue parachute will deploy at approximately 50,000 feet, followed by the main parachute system for touchdown.

Electric solar-wind sail for asteroid touring missions and planetary protection

NASA Astrophysics Data System (ADS)

Janhunen, P.

2014-07-01

The electric solar-wind sail (electric sail, E-sail [1,2]) is a relatively new concept for moving around in the solar system without consuming propellant and by using the thrust provided by the natural solar wind to produce propulsion. The E-sail is based on deploying, using the centrifugal force, a set of long, thin metallic tethers and charging them to high positive voltage by actively removing negative charge from the system by an electron gun. To make the tethers resistant towards inevitable wire cuts by micrometeoroids, they must be made by bonding from multiple (typically 4) thin (25--50 μ m) aluminium wires. Production of the tethers was a technical challenge which was recently overcome. According to present numerical estimates, the E-sail could produce up to 1 N of propellantless thrust out of less than 200 kg package which is enough to give characteristic acceleration of 1 mm/s^2 to a spacecraft weighing 1 tonne, thus producing 30 km/s of delta-v per year. The thrust scales as ˜ 1/r where r is the solar distance. There are ways to control and vector the thrust enough to enable inward and outward spiralling missions in the solar system. The E-sail working principle has been indirectly measured in a laboratory, and ESTCube-1 CubeSat experiment is underway in orbit (in late March 2014 it was waiting to be started) to measure the E-sail thrust acting on a short 10-m long tether. A full-scale mission requires ˜ 1000 km of tether altogether (weighing ˜10 kg). The production of a 1-km piece of tether has been demonstrated in laboratory [3]. If the E-sail holds up its present promise, it would be ideally suited for asteroid missions because it enables production of similar level of thrust than ion engines, but needs only a small fraction of the electric power and never runs out of propellant because it does not use any (the ''propellant'' being the natural solar-wind plasma flow). Here we consider especially a mission which would tour the asteroid belt for a long time, moving from asteroid to asteroid in a bit similar way as, e.g., Mars rovers move from rock to rock on the planet's surface. After starting from the Earth, the mission would slowly spiral outward, making rendezvous with interesting asteroids along the way, as well as flybys or even a larger number of asteroids as opportunities arise. The spacecraft would do remote sensing of the bodies and perhaps also deploy small CubeSat-sized expendable landers on them (the mother spacecraft cannot land on an asteroid or else it would lose the E-sail tethers). The mission would first explore near-Earth objects, then pass through the main belt and end up with the Trojans, exploring asteroids in rendezvous and flyby modes all the time. Asteroids in roughly circular orbits and at low inclination would be the easiest and most likely targets for rendezvous mode encounters, while there would be less restrictions for flyby mode observations. Besides for pure asteroid science, the E-sail could also be used for planetary protection, either through direct propulsive deflection of a dangerous asteroid [4] or by accelerating a relatively lightweight impactor spacecraft to a retrograde orbit and in that way maximizing the available deflecting impact energy for given impactor mass. E-sails could take a number of such impactors to retrograde storage orbits from which they could be commanded to impact a dangerous asteroid with relatively short warning time. Such impactor fleet would not be dangerous to the Earth because the vehicles can be designed to burn completely in the atmosphere, in the unlikely event that due to some mishap one of them would collide with the Earth. The E-sail has potentially large applicability to asteroids as it promises ''free'' transportation in the solar system. As a next step, a solar-wind test mission is needed to demonstrate the technology in the authentic environment.

Population trends of binary near-Earth asteroids based on radar and lightcurves observations

NASA Astrophysics Data System (ADS)

Brozovic, Marina; Benner, Lance A. M.; Naidu, Shantanu P.; Taylor, Patrick A.; Busch, Michael W.; Margot, Jean-Luc; Nolan, Michael C.; Howell, Ellen S.; Springmann, Alessondra; Giorgini, Jon D.; Shepard, Michael K.; Magri, Christopher; Richardson, James E.; Rivera-Valentin, Edgard G.; Rodriguez-Ford, Linda A.; Zambrano Marin, Luisa Fernanda

2016-10-01

The Arecibo and Goldstone planetary radars are invaluable instruments for the discovery and characterization of binary and triple asteroids in the near-Earth asteroid (NEA) population. To date, 41 out of 56 known binaries and triples (~73% of the objects) have been discovered by radar and 49 of these multiple systems have been detected by radar. Their absolute magnitudes range from 12.4 for (1866) Sisyphus to 22.6 for 2015 TD144 and have a mean and rms dispersion of 18.1+-2.0. There is a pronounced decrease in the abundance of binaries for absolute magnitudes H>20. One of the smallest binaries, 1994 CJ1, with an absolute magnitude H=21.4, is also the most accessible binary for a spacecraft rendezvous. Among 365 NEAs with H<22 (corresponding to diameters larger than ~ 140 m) detected by radar since 1999, ~13% have at least one companion. Two triple systems are known, (15391) 2001 SN263 and (136617) 1994 CC, but this is probably an underestimate due to low signal to noise ratios (SNRs) for many of the binary radar detections. Taxonomic classes have been reported for 41 out of 56 currently known multiple systems and some trends are starting to emerge: at least 50% of multiple asteroid systems are S, Sq, Q, or Sk, and at least 20% are optically dark (C, B, P, or U). Thirteen V-class NEAs have been observed by radar and six of them are binaries. Curiously, a comparable number of E-class objects have been detected by radar, but none is known to be a binary.

International Space Station (ISS)

NASA Image and Video Library

2005-07-28

Launched on July 26 2005 from the Kennedy Space Center in Florida, STS-114 was classified as Logistics Flight 1. Among the Station-related activities of the mission were the delivery of new supplies and the replacement of one of the orbital outpost's Control Moment Gyroscopes (CMGs). STS-114 also carried the Raffaello Multi-Purpose Logistics Module (MPLM) and the External Stowage Platform-2. Back dropped by popcorn-like clouds, the MPLM can be seen in the cargo bay as Discovery undergoes rendezvous and docking operations. Cosmonaut Sergei K. Kriklev, Expedition 11 Commander, and John L. Phillips, NASA Space Station officer and flight engineer photographed the spacecraft from the International Space Station (ISS).

International Space Station (ISS)

NASA Image and Video Library

2005-07-28

Launched on July 26, 2005 from the Kennedy Space Center in Florida, STS-114 was classified as Logistics Flight 1. Among the Station-related activities of the mission were the delivery of new supplies and the replacement of one of the orbital outpost's Control Moment Gyroscopes (CMGs). STS-114 also carried the Raffaello Multi-Purpose Logistics Module (MPLM) and the External Stowage Platform-2. Back dropped by popcorn-like clouds, the MPLM can be seen in the cargo bay as Discovery undergoes rendezvous and docking operations. Cosmonaut Sergei K. Kriklev, Expedition 11 Commander, and John L. Phillips, NASA Space Station officer and flight engineer photographed the spacecraft from the International Space Station (ISS).

KSC-99pc45

NASA Image and Video Library

1999-01-11

Bright white light (left) and blue light (upper right) appear on the solar panels of the Stardust spacecraft during lighting tests in the Payload Hazardous Servicing Facility. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule to be jettisoned as it swings by Earth in January 2006

SEP thrust subsystem performance sensitivity analysis

NASA Technical Reports Server (NTRS)

Atkins, K. L.; Sauer, C. G., Jr.; Kerrisk, D. J.

1973-01-01

This is a two-part report on solar electric propulsion (SEP) performance sensitivity analysis. The first part describes the preliminary analysis of the SEP thrust system performance for an Encke rendezvous mission. A detailed description of thrust subsystem hardware tolerances on mission performance is included together with nominal spacecraft parameters based on these tolerances. The second part describes the method of analysis and graphical techniques used in generating the data for Part 1. Included is a description of both the trajectory program used and the additional software developed for this analysis. Part 2 also includes a comprehensive description of the use of the graphical techniques employed in this performance analysis.

KSC-99pc48

NASA Image and Video Library

1999-01-11

In the Payload Hazardous Servicing Facility, workers get ready to rotate the Stardust spacecraft before deploying the solar panels (at left and right) for lighting tests. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule to be jettisoned as it swings by Earth in January 2006

KSC-99pc47

NASA Image and Video Library

1999-01-11

In the Payload Hazardous Servicing Facility, workers raise the Stardust spacecraft from its workstand to move it to another area for lighting tests on the solar panels. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule to be jettisoned as it swings by Earth in January 2006

KSC-99pc44

NASA Image and Video Library

1999-01-11

In the Payload Hazardous Servicing Facility, a worker looks over the solar panels of the Stardust spacecraft before it undergoes lighting tests. Stardust is scheduled to be launched aboard a Boeing Delta II rocket from Launch Pad 17A, Cape Canaveral Air Station, on Feb. 6, 1999, for a rendezvous with the comet Wild 2 in January 2004. Stardust will use a substance called aerogel to capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a sample return capsule (its white cap is seen on the left) to be jettisoned as it swings by Earth in January 2006

Space shuttle Atlantis preparing to dock with Mir space station

NASA Image and Video Library

1995-06-28

NM18-309-018 (28 June 1995) --- The Space Shuttle Atlantis orbits Earth at a point above Iraq as photographed by one of the Mir-18 crew members aboard Russia's Mir Space Station. The image was photographed prior to rendezvous and docking of the two spacecraft. The Spacelab science module and the tunnel connecting it to the crew cabin, as well as the added mechanism for interface with the Mir's docking system can be easily seen. The geography pictured is 60 miles northwest of Baghdad. The Buhayrat Ath Tharthar (reservoir) is the widest body of water visible. Also seen are the Tigris and Euphrates Rivers.

Tethers

NASA Technical Reports Server (NTRS)

Cutler, Andrew Hall; Carroll, Joseph A.

1992-01-01

A tether of sufficient strength, capable of being lengthened or shortened and having appropriate apparatuses for capturing and releasing bodies at its ends, may be useful in propulsion applications. For example, a tether could allow rendezvous between spacecraft in substantially different orbits without using propellant. A tether could also allow co-orbiting spacecraft to exchange momentum and separate. Thus, a reentering spacecraft (such as the Shuttle) could give its momentum to one remaining on orbit (such as the space station). Similarly, a tether facility could gain momentum from a high I(sub sp)/low thrust mechanism (which could be an electrodynamics tether) and transfer than momentum by means of a tether to payloads headed for many different orbits. Such a facility would, in effect, combine high I(sub sp) with high thrust, although only briefly. An electrodynamic tether could propel a satellite from its launch inclination to a higher or lower inclination. Tethers could also allow samples to be taken from bodies such as the Moon. Three types of tether operations are illustrated. The following topics are discussed: (1) tether characteristics; (2) tether propulsion methods--basics, via momentum transfer, and electrodynamic tether propulsion; and (3) their use in planetary exploration.

KSC-108-75P-0005

NASA Image and Video Library

1975-01-14

CAPE CANAVERAL, Fla. – Model of docked Apollo and Soyuz spacecraft in the foreground and skylight in the Vehicle Assembly Building high bay frame the second stage of the Saturn 1B booster that will launch the United States ASTP mission as a crane raises it prior to its mating with the Saturn 1B first stage. Mating of the Saturn 1B first and second stages was completed this morning. The U. S. ASTP launch with mission commander Thomas Stafford, command module pilot Vance Brand and docking module pilot Donald Slayton is scheduled at 3:50 p.m. EDT July 15. The first international crewed spaceflight was a joint U.S.-U.S.S.R. rendezvous and docking mission. The Apollo-Soyuz Test Project, or ASTP, took its name from the spacecraft employed: the American Apollo and the Soviet Soyuz. The three-man Apollo crew lifted off from Kennedy Space Center aboard a Saturn IB rocket on July 15, 1975, to link up with the Soyuz that had launched a few hours earlier. A cylindrical docking module served as an airlock between the two spacecraft for transfer of the crew members. Photo credit: NASA

Performance modelling of miniaturized flash-imaging lidars for future mars exploration missions

NASA Astrophysics Data System (ADS)

Mitev, V.; Pollini, A.; Haesler, J.; Pereira do Carmo, João.

2017-11-01

Future planetary exploration missions require the support of 3D vision in the GN&C during key spacecraft's proximity phases, namely: i) spacecraft precision and soft Landing on the planet's surface; ii) Rendezvous and Docking (RVD) between a Sample Canister (SC) and an orbiter spacecraft; iii) Rover Navigation (RN) on planetary surface. The imaging LiDARs are among the best candidate for such tasks [1-3]. The combination of measurement requirements and environmental conditions seems to find its optimum in the flash 3D LiDAR architecture. Here we present key steps is the evaluation of novelty light detectors and MOEMS (Micro-Opto- Electro-Mechanical Systems) technologies with respect to LiDAR system performance and miniaturization. The objectives of the project MILS (Miniaturized Imaging LiDAR System, Phase 1) concentrated on the evaluation of novel detection and scanning technologies for the miniaturization of 3D LiDARs intended for planetary mission. Preliminary designs for an elegant breadboard (EBB) for the three tasks stated above (Landing, RVD and RN) were proposed, based on results obtained with a numerical model developed in the project and providing the performances evaluation of imaging LiDARs.

Fast, Safe, Propellant-Efficient Spacecraft Motion Planning Under Clohessy-Wiltshire-Hill Dynamics

NASA Technical Reports Server (NTRS)

Starek, Joseph A.; Schmerling, Edward; Maher, Gabriel D.; Barbee, Brent W.; Pavone, Marco

2016-01-01

This paper presents a sampling-based motion planning algorithm for real-time and propellant-optimized autonomous spacecraft trajectory generation in near-circular orbits. Specifically, this paper leverages recent algorithmic advances in the field of robot motion planning to the problem of impulsively actuated, propellant- optimized rendezvous and proximity operations under the Clohessy-Wiltshire-Hill dynamics model. The approach calls upon a modified version of the FMT* algorithm to grow a set of feasible trajectories over a deterministic, low-dispersion set of sample points covering the free state space. To enforce safety, the tree is only grown over the subset of actively safe samples, from which there exists a feasible one-burn collision-avoidance maneuver that can safely circularize the spacecraft orbit along its coasting arc under a given set of potential thruster failures. Key features of the proposed algorithm include 1) theoretical guarantees in terms of trajectory safety and performance, 2) amenability to real-time implementation, and 3) generality, in the sense that a large class of constraints can be handled directly. As a result, the proposed algorithm offers the potential for widespread application, ranging from on-orbit satellite servicing to orbital debris removal and autonomous inspection missions.

Flag and Footprints Mission Mars: Preliminary Design Review Two

NASA Astrophysics Data System (ADS)

1998-01-01

SMI has developed a preliminary guideline for a flag and footprints manned mission to Mars. The manned mission is a split mission where the return and ground supplies will be sent on a cargo spacecraft. The crew spacecraft will leave on a high-energy trajectory once the cargo spacecraft has arrived in the prescribed orbit about Mars. The trajectory will be approximately 150-day from Low Earth Orbit (LEO) to the prescribed rendezvous orbit. The crew spacecraft will then dock with the orbiting cargo spacecraft for refuel and resupply. In addition, once safely docked, the crew members will transfer to the Mars Excursion Vehicle (MEV) for transport to the Martian surface. Each vehicle will be equipped with all necessary subsystems. To facilitate the transport of a large payload from Earth to Mars, the cargo spacecraft will utilize Ion propulsion. The Ion propulsion is ideal due to the high Isp characteristics. The crew spacecraft will be propelled with high-thrust RL-10 engines. Due to the smaller mass of the crew spacecraft, the spacecraft will utilize a 150-day high-energy trajectory. The MEV propulsion will be hypergolic. This choice of fuel is due to the reliability and simplicity of use. The crew members will stay on the surface of Mars for 30-days. During the 30-days, the crew will perform a series of scientific and exploratory experiments. To broaden the astronauts range of exploration, the astronauts will have access to three Unmanned Aerial Vehicles (UAV) and one rover while on the surface. The scientific experiments will consist of several soil and rock analyses as well as atmospheric study. Upon completion of the 30-day ground phase, the astronauts will return to the orbiting crew ship for return to Earth. SMI's flag and footprints mission outlines the fundamental systems and general requirements for these systems. SMI feels that with the fulfillment of these fundamental systems, this mission will be a highly desirable and potential candidate for development by NASA.

Saturn Apollo Program

NASA Image and Video Library

1964-09-09

This is the official portrait of astronaut Frank Borman. A career Air Force officer from 1950, his assignments included service as a fighter pilot, an operational pilot and instructor, an experimental test pilot and an assistant professor of thermodynamics and fluid mechanics at West Point. When selected by NASA, Frank Borman was an instructor at the Aerospace Research Pilot School at Edwards AFB, California. In 1967 he served as a member of the Apollo 204 Fire Investigation Board, investigating the causes of the fire which killed three astronauts aboard an Apollo spacecraft. Later he became the Apollo Program Resident Manager, heading the team that reengineered the Apollo spacecraft. He also served as Field Director of the NASA Space Station Task Force. Frank Borman retired from the air Force in 1970, but is well remembered as a part of American history as a pioneer in the exploration of space. He is a veteran of both the Gemini 7, 1965 Space Orbital Rendezvous with Gemini 6 and the first manned lunar orbital mission, Apollo 8, in 1968.

Comet nucleus sample return mission

NASA Technical Reports Server (NTRS)

1983-01-01

A comet nucleus sample return mission in terms of its relevant science objectives, candidate mission concepts, key design/technology requirements, and programmatic issues is discussed. The primary objective was to collect a sample of undisturbed comet material from beneath the surface of an active comet and to preserve its chemical and, if possible, its physical integrity and return it to Earth in a minimally altered state. The secondary objectives are to: (1) characterize the comet to a level consistent with a rendezvous mission; (2) monitor the comet dynamics through perihelion and aphelion with a long lived lander; and (3) determine the subsurface properties of the nucleus in an area local to the sampled core. A set of candidate comets is discussed. The hazards which the spacecraft would encounter in the vicinity of the comet are also discussed. The encounter strategy, the sampling hardware, the thermal control of the pristine comet material during the return to Earth, and the flight performance of various spacecraft systems and the cost estimates of such a mission are presented.

Spacecraft Chemical Propulsion Systems at NASA's Marshall Space Flight Center: Heritage and Capabilities

NASA Technical Reports Server (NTRS)

McRight, Patrick S.; Sheehy, Jeffrey A.; Blevins, John A.

2005-01-01

NASA Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V and the Space Shuttle. This paper highlights a lesser known but equally rich side of MSFC - its heritage in spacecraft chemical propulsion systems and its current capabilities for in-space propulsion system development and chemical propulsion research. The historical narrative describes the efforts associated with developing upper-stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology, X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and several technology development activities. Also discussed are MSFC chemical propulsion research capabilities, along with near- and long-term technology challenges to which MSFC research and system development competencies are relevant.

Comparison of candidate solar array maximum power utilization approaches. [for spacecraft propulsion

NASA Technical Reports Server (NTRS)

Costogue, E. N.; Lindena, S.

1976-01-01

A study was made of five potential approaches that can be utilized to detect the maximum power point of a solar array while sustaining operations at or near maximum power and without endangering stability or causing array voltage collapse. The approaches studied included: (1) dynamic impedance comparator, (2) reference array measurement, (3) onset of solar array voltage collapse detection, (4) parallel tracker, and (5) direct measurement. The study analyzed the feasibility and adaptability of these approaches to a future solar electric propulsion (SEP) mission, and, specifically, to a comet rendezvous mission. Such missions presented the most challenging requirements to a spacecraft power subsystem in terms of power management over large solar intensity ranges of 1.0 to 3.5 AU. The dynamic impedance approach was found to have the highest figure of merit, and the reference array approach followed closely behind. The results are applicable to terrestrial solar power systems as well as to other than SEP space missions.

KSC-05PD-0126

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. On Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., shadows paint the Boeing Delta II rocket carrying the Deep Impact spacecraft as the mobile service tower at left is rolled back before launch.Scheduled for liftoff at 1:47 p.m. EST today, Deep Impact will head for space and a rendezvous with Comet Tempel 1 when the comet is 83 million miles from Earth. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network. Deep Impact is a NASA Discovery mission.

KSC-05PD-0125

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. On Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., the Boeing Delta II rocket carrying the Deep Impact spacecraft looms into the night sky as the mobile service tower at right is rolled back before launch. Scheduled for liftoff at 1:47 p.m. EST today, Deep Impact will head for space and a rendezvous with Comet Tempel 1 when the comet is 83 million miles from Earth. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network. Deep Impact is a NASA Discovery mission.

KSC-05PD-0127

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. On Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., the Boeing Delta II carrying the Deep Impact spacecraft rocket shines under spotlights in the early dawn hours as it waits for launch. Scheduled for liftoff at 1:47 p.m. EST today, Deep Impact will head for space and a rendezvous with Comet Tempel 1 when the comet is 83 million miles from Earth. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network. Deep Impact is a NASA Discovery mission.

KSC-05PD-0129

NASA Technical Reports Server (NTRS)

2005-01-01

KENNEDY SPACE CENTER, FLA. The sun rises behind Launch Pad 17-B, Cape Canaveral Air Force Station, Fla., where the Boeing Delta II rocket carrying the Deep Impact spacecraft waits for launch. Gray clouds above the horizon belie the favorable weather forecast for the afternoon launch. Scheduled for liftoff at 1:47 p.m. EST today, Deep Impact will head for space and a rendezvous with Comet Tempel 1 when the comet is 83 million miles from Earth. After releasing a 3- by 3-foot projectile (impactor) to crash onto the surface July 4, 2005, Deep Impacts flyby spacecraft will reveal the secrets of the comets interior by collecting pictures and data of how the crater forms, measuring the craters depth and diameter as well as the composition of the interior of the crater and any material thrown out, and determining the changes in natural outgassing produced by the impact. It will send the data back to Earth through the antennas of the Deep Space Network. Deep Impact is a NASA Discovery mission.

Design concepts and performance of NASA X-band (7162 MHz/8415 MHz) transponder for deep-space spacecraft applications

NASA Technical Reports Server (NTRS)

Mysoor, N. R.; Perret, J. D.; Kermode, A. W.

1991-01-01

The design concepts and measured performance characteristics are summarized of an X band (7162 MHz/8415 MHz) breadboard deep space transponder (DSP) for future spacecraft applications, with the first use scheduled for the Comet Rendezvous Asteroid Flyby (CRAF) and Cassini missions in 1995 and 1996, respectively. The DST consists of a double conversion, superheterodyne, automatic phase tracking receiver, and an X band (8415 MHz) exciter to drive redundant downlink power amplifiers. The receiver acquires and coherently phase tracks the modulated or unmodulated X band (7162 MHz) uplink carrier signal. The exciter phase modulates the X band (8415 MHz) downlink signal with composite telemetry and ranging signals. The receiver measured tracking threshold, automatic gain control, static phase error, and phase jitter characteristics of the breadboard DST are in good agreement with the expected performance. The measured results show a receiver tracking threshold of -158 dBm and a dynamic signal range of 88 dB.

Economic and technical aspects of repair, servicing, and retrieval of low earth orbit free flying spacecraft

NASA Technical Reports Server (NTRS)

Cepollina, F. J.

1982-01-01

The economic and technical aspects of the Solar Maximum Observatory Repair Mission at NASA are presented, in an effort to demonstrate the Space Shuttle capability to rendezvous with and repair on-orbit the Solar Maximum Observatory (SMM). A failure in the Attitude Control Subsystem (ACS) after 10 months of operation caused a loss in precision pointing capability. The Multimission Modular Spacecraft (MMS) used for the mission, was designed with on-orbit repairability, and to correct various instrument anomalies, repiar kits such as an electronics box, a thermal aperture closure, and a high energy particle reflection baffle will be used. In addition, a flight support system will be used to berth, electrically safe, and support all the repair activities. A two year effort is foreseen, and the economic return on SMM will be $176 M, in addition to two to three years of solar observation. The mission will eventually conduct studies on flare as a function of solar cycle.

The U.S. Rosetta Project: Preparations for Prime Mission, 2014

NASA Technical Reports Server (NTRS)

Alexander, C.; Chmielewski, A.; Aguinaldo, A. M.; Ko, A.; Accomazzo, A.; Taylor, M. G. G.

2014-01-01

In 2014, the International Rosetta mission will place a spacecraft in orbit around comet 67P/Churyumov-Gerasimenko and deliver a lander to the comet's surface. The National Aeronautics and Space Administration's (NASA) contribution to the International Rosetta mission, designated the U.S. Rosetta Project, includes several instruments, tracking support, and science support for some non-US payloads. In July 2011 the spacecraft was placed in a long-duration hibernation mode planned to last approximately 37 months to conserve electrical power. Rosetta will rendezvous with 67P/Churyumov-Gerasimenko in 2014. On the eve of the mission's arrival at its target, this paper highlights three issues related to Rosetta's looming prime mission: (A) measures taken in 2009 to prepare the US Rosetta Project for the long-duration hibernation mode; (B) risk reviews conducted in 2013 to prepare the US Rosetta Project for exit from hibernation; (C) ESA and NASA preparations for use of NASA Deep Space Network (DSN) assets related to keyword files.

KSC-75PC-0330

NASA Image and Video Library

1975-07-03

CAPE CANAVERAL, Fla. – ASTP prime crewmen Donald Slayton, Thomas Stafford and Vance Brand pose with their Saturn IB launch vehicle following the Countdown Demonstration Test [CDDT], a step-by-step dress rehearsal for their July 15 launch. During the “wet” portion of the test, conducted yesterday, the stages of the launch vehicle were fueled as they will be on launch day. The fuels were off loaded and the terminal portion of the count repeated today with the astronauts aboard the vehicle. The first international crewed spaceflight was a joint U.S.-U.S.S.R. rendezvous and docking mission. The Apollo-Soyuz Test Project, or ASTP, took its name from the spacecraft employed: the American Apollo and the Soviet Soyuz. The three-man Apollo crew lifted off from Kennedy Space Center aboard a Saturn IB rocket on July 15, 1975, to link up with the Soyuz that had launched a few hours earlier. A cylindrical docking module served as an airlock between the two spacecraft for transfer of the crew members. Photo credit: NASA

The Data Transfer Kit: A geometric rendezvous-based tool for multiphysics data transfer

DOE Office of Scientific and Technical Information (OSTI.GOV)

Slattery, S. R.; Wilson, P. P. H.; Pawlowski, R. P.

2013-07-01

The Data Transfer Kit (DTK) is a software library designed to provide parallel data transfer services for arbitrary physics components based on the concept of geometric rendezvous. The rendezvous algorithm provides a means to geometrically correlate two geometric domains that may be arbitrarily decomposed in a parallel simulation. By repartitioning both domains such that they have the same geometric domain on each parallel process, efficient and load balanced search operations and data transfer can be performed at a desirable algorithmic time complexity with low communication overhead relative to other types of mapping algorithms. With the increased development efforts in multiphysicsmore » simulation and other multiple mesh and geometry problems, generating parallel topology maps for transferring fields and other data between geometric domains is a common operation. The algorithms used to generate parallel topology maps based on the concept of geometric rendezvous as implemented in DTK are described with an example using a conjugate heat transfer calculation and thermal coupling with a neutronics code. In addition, we provide the results of initial scaling studies performed on the Jaguar Cray XK6 system at Oak Ridge National Laboratory for a worse-case-scenario problem in terms of algorithmic complexity that shows good scaling on 0(1 x 104) cores for topology map generation and excellent scaling on 0(1 x 105) cores for the data transfer operation with meshes of O(1 x 109) elements. (authors)« less

Overview of the Multi-Spectral Imager on the NEAR spacecraft

NASA Astrophysics Data System (ADS)

Hawkins, S. E., III

1996-07-01

The Multi-Spectral Imager on the Near Earth Asteroid Rendezvous (NEAR) spacecraft is a 1 Hz frame rate CCD camera sensitive in the visible and near infrared bands (~400-1100 nm). MSI is the primary instrument on the spacecraft to determine morphology and composition of the surface of asteroid 433 Eros. In addition, the camera will be used to assist in navigation to the asteroid. The instrument uses refractive optics and has an eight position spectral filter wheel to select different wavelength bands. The MSI optical focal length of 168 mm gives a 2.9 ° × 2.25 ° field of view. The CCD is passively cooled and the 537×244 pixel array output is digitized to 12 bits. Electronic shuttering increases the effective dynamic range of the instrument by more than a factor of 100. A one-time deployable cover protects the instrument during ground testing operations and launch. A reduced aperture viewport permits full field of view imaging while the cover is in place. A Data Processing Unit (DPU) provides the digital interface between the spacecraft and the Camera Head and uses an RTX2010 processor. The DPU provides an eight frame image buffer, lossy and lossless data compression routines, and automatic exposure control. An overview of the instrument is presented and design parameters and trade-offs are discussed.

Two-phase framework for near-optimal multi-target Lambert rendezvous

NASA Astrophysics Data System (ADS)

Bang, Jun; Ahn, Jaemyung

2018-03-01

This paper proposes a two-phase framework to obtain a near-optimal solution of multi-target Lambert rendezvous problem. The objective of the problem is to determine the minimum-cost rendezvous sequence and trajectories to visit a given set of targets within a maximum mission duration. The first phase solves a series of single-target rendezvous problems for all departure-arrival object pairs to generate the elementary solutions, which provides candidate rendezvous trajectories. The second phase formulates a variant of traveling salesman problem (TSP) using the elementary solutions prepared in the first phase and determines the final rendezvous sequence and trajectories of the multi-target rendezvous problem. The validity of the proposed optimization framework is demonstrated through an asteroid exploration case study.

Automated target recognition and tracking using an optical pattern recognition neural network

NASA Technical Reports Server (NTRS)

Chao, Tien-Hsin

1991-01-01

The on-going development of an automatic target recognition and tracking system at the Jet Propulsion Laboratory is presented. This system is an optical pattern recognition neural network (OPRNN) that is an integration of an innovative optical parallel processor and a feature extraction based neural net training algorithm. The parallel optical processor provides high speed and vast parallelism as well as full shift invariance. The neural network algorithm enables simultaneous discrimination of multiple noisy targets in spite of their scales, rotations, perspectives, and various deformations. This fully developed OPRNN system can be effectively utilized for the automated spacecraft recognition and tracking that will lead to success in the Automated Rendezvous and Capture (AR&C) of the unmanned Cargo Transfer Vehicle (CTV). One of the most powerful optical parallel processors for automatic target recognition is the multichannel correlator. With the inherent advantages of parallel processing capability and shift invariance, multiple objects can be simultaneously recognized and tracked using this multichannel correlator. This target tracking capability can be greatly enhanced by utilizing a powerful feature extraction based neural network training algorithm such as the neocognitron. The OPRNN, currently under investigation at JPL, is constructed with an optical multichannel correlator where holographic filters have been prepared using the neocognitron training algorithm. The computation speed of the neocognitron-type OPRNN is up to 10(exp 14) analog connections/sec that enabling the OPRNN to outperform its state-of-the-art electronics counterpart by at least two orders of magnitude.

Improving spacecraft design using a multidisciplinary design optimization methodology

NASA Astrophysics Data System (ADS)

Mosher, Todd Jon

2000-10-01

Spacecraft design has gone from maximizing performance under technology constraints to minimizing cost under performance constraints. This is characteristic of the "faster, better, cheaper" movement that has emerged within NASA. Currently spacecraft are "optimized" manually through a tool-assisted evaluation of a limited set of design alternatives. With this approach there is no guarantee that a systems-level focus will be taken and "feasibility" rather than "optimality" is commonly all that is achieved. To improve spacecraft design in the "faster, better, cheaper" era, a new approach using multidisciplinary design optimization (MDO) is proposed. Using MDO methods brings structure to conceptual spacecraft design by casting a spacecraft design problem into an optimization framework. Then, through the construction of a model that captures design and cost, this approach facilitates a quicker and more straightforward option synthesis. The final step is to automatically search the design space. As computer processor speed continues to increase, enumeration of all combinations, while not elegant, is one method that is straightforward to perform. As an alternative to enumeration, genetic algorithms are used and find solutions by reviewing fewer possible solutions with some limitations. Both methods increase the likelihood of finding an optimal design, or at least the most promising area of the design space. This spacecraft design methodology using MDO is demonstrated on three examples. A retrospective test for validation is performed using the Near Earth Asteroid Rendezvous (NEAR) spacecraft design. For the second example, the premise that aerobraking was needed to minimize mission cost and was mission enabling for the Mars Global Surveyor (MGS) mission is challenged. While one might expect no feasible design space for an MGS without aerobraking mission, a counterintuitive result is discovered. Several design options that don't use aerobraking are feasible and cost effective. The third example is an original commercial lunar mission entitled Eagle-eye. This example shows how an MDO approach is applied to an original mission with a larger feasible design space. It also incorporates a simplified business case analysis.

Interactive optimization approach for optimal impulsive rendezvous using primer vector and evolutionary algorithms

NASA Astrophysics Data System (ADS)

Luo, Ya-Zhong; Zhang, Jin; Li, Hai-yang; Tang, Guo-Jin

2010-08-01

In this paper, a new optimization approach combining primer vector theory and evolutionary algorithms for fuel-optimal non-linear impulsive rendezvous is proposed. The optimization approach is designed to seek the optimal number of impulses as well as the optimal impulse vectors. In this optimization approach, adding a midcourse impulse is determined by an interactive method, i.e. observing the primer-magnitude time history. An improved version of simulated annealing is employed to optimize the rendezvous trajectory with the fixed-number of impulses. This interactive approach is evaluated by three test cases: coplanar circle-to-circle rendezvous, same-circle rendezvous and non-coplanar rendezvous. The results show that the interactive approach is effective and efficient in fuel-optimal non-linear rendezvous design. It can guarantee solutions, which satisfy the Lawden's necessary optimality conditions.

Mission options for rendezvous with the most accessible Near-Earth Asteroid - 1989 ML

NASA Technical Reports Server (NTRS)

Mcadams, Jim V.

1992-01-01

The recent discovery of the Amor-class 1989 ML, the most accessible known asteroid for minimum-energy rendezvous missions, has expedited the search for frequent, low-cost Near-Earth Asteroid rendezvous and round-trip missions. This paper identifies trajectory characteristics and assesses mass performance for low Delta V ballistic rendezvous opportunities to 1989 ML during the period 1996-2010. This asteroid also offers occasional unique extended mission opportunities, such as the lowest known Delta V requirement for any asteroid sample return mission as well as pre-rendezvous asteroid flyby and post-rendezvous comet flyby opportunities requiring less than 5.25 km/sec total Delta V. This paper also briefly comments concerning mission opportunities for asteroid 1991 JW, which recently replaced other known asteroids as the most accessible Near-Earth Asteroid for fast rendezvous and round-trip missions.

Synchronized Position Hold, Engage, Reorient, Experimental Satellites

NASA Technical Reports Server (NTRS)

Miller, David W.; Wilson, Edward; How, Jonathan; Sanenz-Otero, Alvar; Chamitoff, Gregory

2009-01-01

Synchronized Position Hold, Engage, Reorient, Experimental Satellites (SPHERES) are bowling-ball sized spherical satellites. They will be used inside the space station to test a set of well-defined instructions for spacecraft performing autonomous rendezvous and docking maneuvers. Three free-flying spheres will fly within the cabin of the station, performing flight formations. Each satellite is self-contained with power, propulsion, computers and navigation equipment. The results are important for satellite servicing, vehicle assembly and formation flying spacecraft configurations. SPHERES is a testbed for formation flying by satellites, the theories and calculations that coordinate the motion of multiple bodies maneuvering in microgravity. To achieve this inside the ISS cabin, bowling-ball-sized spheres perform various maneuvers (or protocols), with one to three spheres operating simultaneously . The Synchronized Position Hold, Engage, Reorient, Experimental Satellites (SPHERES) experiment will test relative attitude control and station-keeping between satellites, re-targeting and image plane filling maneuvers, collision avoidance and fuel balancing algorithms, and an array of geometry estimators used in various missions. SPHERES consists of three self-contained satellites, which are 18 sided polyhedrons that are 0.2 meter in diameter and weigh 3.5 kilograms. Each satellite contains an internal propulsion system, power, avionics, software, communications, and metrology subsystems. The propulsion system uses CO2, which is expelled through the thrusters. SPHERES satellites are powered by AA batteries. The metrology subsystem provides real-time position and attitude information. To simulate ground station-keeping, a laptop will be used to transmit navigational data and formation flying algorithms. Once these data are uploaded, the satellites will perform autonomously and hold the formation until a new command is given.

The Ion Propulsion System on NASA's Space Technology 4/Champollion Comet Rendezvous Mission

NASA Technical Reports Server (NTRS)

Brophy, John R.; Garner, Charles E.; Weiss, Jeffery M.

1999-01-01

The ST4/Champollion mission is designed to rendezvous with and land on the comet Tempel 1 and return data from the first-ever sampling of a comet surface. Ion propulsion is an enabling technology for this mission. The ion propulsion system on ST4 consists of three ion engines each essentially identical to the single engine that flew on the DS1 spacecraft. The ST4 propulsion system will operate at a maximum input power of 7.5 kW (3.4 times greater than that demonstrated on DS1), will produce a maximum thrust of 276 mN, and will provide a total (Delta)V of 11.4 km/s. To accomplish this the propulsion system will carry 385 kg of xenon. All three engines will be operated simultaneously for the first 168 days of the mission. The nominal mission requires that each engine be capable of processing 118 kg. If one engine fails after 168 days, the remaining two engines can perform the mission, but must be capable of processing 160 kg of xenon, or twice the original thruster design requirement. Detailed analyses of the thruster wear-out failure modes coupled with experience from long-duration engine tests indicate that the thrusters have a high probability of meeting the 160-kg throughput requirement.

Orbital Express Advanced Video Guidance Sensor

NASA Technical Reports Server (NTRS)

Howard, Ricky; Heaton, Andy; Pinson, Robin; Carrington, Connie

2008-01-01

In May 2007 the first US fully autonomous rendezvous and capture was successfully performed by DARPA's Orbital Express (OE) mission. Since then, the Boeing ASTRO spacecraft and the Ball Aerospace NEXTSat have performed multiple rendezvous and docking maneuvers to demonstrate the technologies needed for satellite servicing. MSFC's Advanced Video Guidance Sensor (AVGS) is a primary near-field proximity operations sensor integrated into ASTRO's Autonomous Rendezvous and Capture Sensor System (ARCSS), which provides relative state knowledge to the ASTRO GN&C system. This paper provides an overview of the AVGS sensor flying on Orbital Express, and a summary of the ground testing and on-orbit performance of the AVGS for OE. The AVGS is a laser-based system that is capable of providing range and bearing at midrange distances and full six degree-of-freedom (6DOF) knowledge at near fields. The sensor fires lasers at two different frequencies to illuminate the Long Range Targets (LRTs) and the Short Range Targets (SRTs) on NEXTSat. Subtraction of one image from the other image removes extraneous light sources and reflections from anything other than the corner cubes on the LRTs and SRTs. This feature has played a significant role for Orbital Express in poor lighting conditions. The very bright spots that remain in the subtracted image are processed by the target recognition algorithms and the inverse-perspective algorithms, to provide 3DOF or 6DOF relative state information. Although Orbital Express has configured the ASTRO ARCSS system to only use AVGS at ranges of 120 m or less, some OE scenarios have provided opportunities for AVGS to acquire and track NEXTSat at greater distances. Orbital Express scenarios to date that have utilized AVGS include a berthing operation performed by the ASTRO robotic arm, sensor checkout maneuvers performed by the ASTRO robotic arm, 10-m unmated operations, 30-m unmated operations, and Scenario 3-1 anomaly recovery. The AVGS performed very well during the pre-unmated operations, effectively tracking beyond its 10-degree Pitch and Yaw limit-specifications, and did not require I-LOAD adjustments before unmated operations. AVGS provided excellent performance in the 10-m unmated operations, effectively tracking and maintaining lock for the duration of this scenario, and showing good agreement between the short and long range targets. During the 30-m unmated operations, the AVGS continuously tracked the SRT to 31.6 m, exceeding expectations, and continuously tracked the LRT from 8.8 m out to 31.6 m, with good agreement between these two target solutions. After this scenario was aborted at a 10-m separation during remate operations, the AVGS tracked the LRT out 54.3 m, until the relative attitude between the vehicles was too large. The vehicles remained apart for eight days, at ranges from 1 km to 6 km. During the approach to remate in this recovery operation, the AVGS began tracking the LRT at 150 m, well beyond the OE planned limits for AVGS ranges, and functioned as the primary sensor for the autonomous rendezvous and docking.

Ion Beam Deflection (AKA Push-Me/Pull-You)

NASA Technical Reports Server (NTRS)

Brophy, John

2013-01-01

The Ion Beam Deflection provides the following potential advantages over other asteroid deflection systems. Like the gravity tractor, it doesn't require despinning of the asteroid. Unlike the gravity tractor, it provides a significantly higher coupling force that is independent of the asteroid size. The concept could be tested as part of the baseline Asteroid Redirect Robotic Mission. The thrust and total impulse are entirely within the design of the SEP vehicle. The total impulse is potentially competitive with kinetic impactors and eliminates the need for a second rendezvous spacecraft.?Gridded ion thrusters provide beam divergence angles of a few degrees enabling long stand-off distances from the asteroid. Mitigating control issues. Minimizing back-sputter contamination risks

View of the Apollo 16 Command/Service Module from the Lunar module in orbit

NASA Image and Video Library

1971-04-20

AS16-113-18282 (23 April 1972) --- The Apollo Command and Service Modules (CSM) "Casper" approaches the Lunar Module (LM) "Orion", from which this photograph was made. The two spacecraft are about to make their final rendezvous of the mission, on April 23, 1972. Astronauts John W. Young, commander, and Charles M. Duke Jr., lunar module pilot, aboard the LM, were returning to the CSM, in lunar orbit, after three successful days on the lunar surface. Astronaut Thomas K. (Ken) Mattingly II, command module pilot, remained with the CSM in lunar orbit, while Young and Duke descended in the LM to explore the Descartes region of the moon.

Raffaello Multi-Purpose Logistics Module (MPLM) in Discovery Cargo Bay

NASA Technical Reports Server (NTRS)

2005-01-01

Launched on July 26, 2005 from the Kennedy Space Center in Florida, STS-114 was classified as Logistics Flight 1. Among the Station-related activities of the mission were the delivery of new supplies and the replacement of one of the orbital outpost's Control Moment Gyroscopes (CMGs). STS-114 also carried the Raffaello Multi-Purpose Logistics Module (MPLM) and the External Stowage Platform-2. Back dropped by popcorn-like clouds, the MPLM can be seen in the cargo bay as Discovery undergoes rendezvous and docking operations. Cosmonaut Sergei K. Kriklev, Expedition 11 Commander, and John L. Phillips, NASA Space Station officer and flight engineer photographed the spacecraft from the International Space Station (ISS).

Raffaello Multi-Purpose Logistics Module (MPLM) in Discovery Cargo Bay

NASA Technical Reports Server (NTRS)

2005-01-01

Launched on July 26 2005 from the Kennedy Space Center in Florida, STS-114 was classified as Logistics Flight 1. Among the Station-related activities of the mission were the delivery of new supplies and the replacement of one of the orbital outpost's Control Moment Gyroscopes (CMGs). STS-114 also carried the Raffaello Multi-Purpose Logistics Module (MPLM) and the External Stowage Platform-2. Back dropped by popcorn-like clouds, the MPLM can be seen in the cargo bay as Discovery undergoes rendezvous and docking operations. Cosmonaut Sergei K. Kriklev, Expedition 11 Commander, and John L. Phillips, NASA Space Station officer and flight engineer photographed the spacecraft from the International Space Station (ISS).

One Year Crew Docking to the International Space Station

NASA Image and Video Library

2015-05-27

This video was taken by the crew members aboard the Soyuz TMA-16M spacecraft which docked to the International Space Station at 9:33 p.m. EDT March 27, 2015. NASA astronaut Scott Kelly and Russian cosmonauts Mikhail Kornienko and Gennady Padalka arrived just six hours after launching from Baikonur, Kazakhstan, completing four orbits around the Earth before catching up with the orbiting laboratory. The vehicle docked to the Poisk module (also known as the Mini-Research Module 2) on the space-facing side of the Russian Service Module. The spinning object in view is an antenna that is part of the automatic rendezvous and docking system known as KURS.

MIT's role in project Apollo. Volume 2: Optical, radar, and candidate subsystems

NASA Technical Reports Server (NTRS)

1972-01-01

The development of optical, radar, and candidate subsystems for Project Apollo is discussed. The design and development of the optical subsystems for both the Apollo command and lunar spacecraft are described. Design approaches, problems, and solutions are presented. The evolution of radar interfaces with the GN&C system is discussed; these interfaces involved both hardware and software in a relatively complex interrelationship. The design and development of three candidate subsystems are also described. The systems were considered for use in Apollo, but were not incorporated into the final GN&C system. The three subsystems discussed are the star tracker-horizon photometer, the map and data viewer and the lunar module optical rendezvous system.

jsc2013e018010

NASA Image and Video Library

2013-03-21

At the Cosmonaut Hotel crew quarters in Baikonur, Kazakhstan, Expedition 35-36 Flight Engineer Chris Cassidy of NASA (left) displays a flight data file book titled “Fast Rendezvous” March 21 as he, Soyuz Commander Pavel Vinogradov (center) and Flight Engineer Alexander Misurkin (right) train for launch to the International Space Station March 29, Kazakh time, in their Soyuz TMA-08M spacecraft from the Baikonur Cosmodrome for a 5 ½ month mission. The “fast rendezvous” refers to the expedited four-orbit, six-hour trip from the launch pad to reach the International Space Station March 29 through an accelerated rendezvous burn plan, the first time this approach will be used for crews flying to the international complex. NASA/Victor Zelentsov

The Rendezvous Monitoring Display Capabilities of the Rendezvous and Proximity Operations Program

NASA Technical Reports Server (NTRS)

Brazzel, Jack; Spehar, Pete; Clark, Fred; Foster, Chris; Eldridge, Erin

2013-01-01

The Rendezvous and Proximity Operations Program (RPOP) is a laptop computer- based relative navigation tool and piloting aid that was developed during the Space Shuttle program. RPOP displays a graphical representation of the relative motion between the target and chaser vehicles in a rendezvous, proximity operations and capture scenario. After being used in over 60 Shuttle rendezvous missions, some of the RPOP display concepts have become recognized as a minimum standard for cockpit displays for monitoring the rendezvous task. To support International Space Station (ISS) based crews in monitoring incoming visiting vehicles, RPOP has been modified to allow crews to compare the Cygnus visiting vehicle s onboard navigated state to processed range measurements from an ISS-based, crew-operated Hand Held Lidar sensor. This paper will discuss the display concepts of RPOP that have proven useful in performing and monitoring rendezvous and proximity operations.

Shuttle OFT Level C navigation requirements

NASA Technical Reports Server (NTRS)

1980-01-01

Detailed requirements for the orbital operations computer loads, OPS 2, and OPS 8 are given. These requirements represent the total on-orbit/rendezvous navigation baseline requirements for the following principal functions: on-orbital/rendezvous navigation sequencer; on-orbit/rendezvous UPP sequencer; on-orbit rendezvous navigation; on-orbit prediction; on-orbit user parameter processing; and landing Site update.

Adaptive lyapunov control and artificial neural networks for spacecraft relative maneuvering using atmospheric differential drag

NASA Astrophysics Data System (ADS)

Perez Chaparro, David Andres

At low Earth orbits, a differential in the drag acceleration between spacecraft can be used to control their relative motion. This drag differential allows for a propellant-free alternative to thrusters for performing relative maneuvers in these orbits. The interest in autonomous propellant-less maneuvering comes from the desire to reduce the costs of spacecraft formations. Formation maneuvering opens up a wide variety of new applications for spacecraft missions, such as on-orbit maintenance and refueling. In this work atmospheric differential drag based nonlinear controllers are presented that can be used for virtually any planar relative maneuver of two spacecraft, provided that there is enough atmospheric density and that the spacecraft can change their ballistic coefficients by sufficient amounts to generate the necessary differential accelerations. The control techniques are successfully tested using high fidelity Satellite Tool Kit simulations for re-phase, fly-around, and rendezvous maneuvers, proving the feasibility of the proposed approach for a real flight. Furthermore, the atmospheric density varies in time and in space as the spacecraft travel along their orbits. The ability to accurately forecast the density allows for accurate onboard orbit propagation and for creating realistic guidance trajectories for maneuvers that rely on the differential drag. In this work a localized density predictor based on artificial neural networks is also presented. The predictor uses density measurements or estimates along the past orbits and can use a set of proxies for solar and geomagnetic activities to predict the value of the density along the future orbits of the spacecraft. The performance of the localized predictor is studied for different neural network structures, testing periods of high and low solar and geomagnetic activities and different prediction windows. Comparison with previously developed methods show substantial benefits in using neural networks, both in prediction accuracy and in the potential for spacecraft onboard implementation. The controllers and the predictor are designed for onboard implementation, and provide spacecraft with the tools necessary for performing propellant-less formation maneuvers using differential drag.

Flex Dynamics Avoidance Control of the NEA Scout Solar Sail Spacecraft's Reaction Control System

NASA Technical Reports Server (NTRS)

Heaton Andrew; Stiltner, Brandon; Diedrich, Benjamin; Becker, Christopher; Orphee, Juan

2017-01-01

The Attitude Control System (ACS) is developed for a Near Earth Asteroid (NEA) Scout mission using a solar sail. The NEA-Scout spacecraft is a 6U cubesat with an 86 square-meter solar sail. NEA Scout will launch on Space Launch System (SLS) Exploration Mission 1 (EM-1), currently scheduled to launch in 2018. The spacecraft will rendezvous with a target asteroid after a two year journey, and will conduct science imagery. The solar sail spacecraft ACS consists of three major actuating subsystems: a Reaction Wheel (RW) control system, a Reaction Control System (RCS), and an Adjustable Mass Translator (AMT) system. The three subsystems allow for a wide range of spacecraft attitude control capabilities, needed for the different phases of the NEA-Scout mission. Because the sail is a flexible structure, care must be taken in designing a control system to avoid exciting the structural modes of the sail. This is especially true for the RCS, which uses pulse actuated, cold-gas jets to control the spacecraft's attitude. While the reaction wheels can be commanded smoothly, the RCS jets are simple on-off actuators. Long duration firing of the RCS jets - firings greater than one second - can be thought of as step inputs to the spacecraft's torque. On the other hand, short duration firings - pulses on the order of 0.1 seconds - can be thought of as impulses in the spacecraft's torque. These types of inputs will excite the structural modes of the spacecraft, causing the sail to oscillate. Sail oscillations are undesirable for many reasons. Mainly, these oscillations will feed into the spacecraft attitude sensors and pointing accuracy, and long term oscillations may be undesirable over the lifetime of the solar sail. In order to limit the sail oscillations, an RCS control scheme is being developed to minimize sail excitations. Specifically, an input shaping scheme similar to the method described in Reference 1 will be employed. A detailed description of the RCS control scheme will be provided with particular emphasis on flexible body excitation. The RCS performance will be provided to show that sail and boom excitation is minimized.

OSIRIS-REx: Sample Return from Asteroid (101955) Bennu

NASA Astrophysics Data System (ADS)

Lauretta, D. S.; Balram-Knutson, S. S.; Beshore, E.; Boynton, W. V.; Drouet d'Aubigny, C.; DellaGiustina, D. N.; Enos, H. L.; Golish, D. R.; Hergenrother, C. W.; Howell, E. S.; Bennett, C. A.; Morton, E. T.; Nolan, M. C.; Rizk, B.; Roper, H. L.; Bartels, A. E.; Bos, B. J.; Dworkin, J. P.; Highsmith, D. E.; Lorenz, D. A.; Lim, L. F.; Mink, R.; Moreau, M. C.; Nuth, J. A.; Reuter, D. C.; Simon, A. A.; Bierhaus, E. B.; Bryan, B. H.; Ballouz, R.; Barnouin, O. S.; Binzel, R. P.; Bottke, W. F.; Hamilton, V. E.; Walsh, K. J.; Chesley, S. R.; Christensen, P. R.; Clark, B. E.; Connolly, H. C.; Crombie, M. K.; Daly, M. G.; Emery, J. P.; McCoy, T. J.; McMahon, J. W.; Scheeres, D. J.; Messenger, S.; Nakamura-Messenger, K.; Righter, K.; Sandford, S. A.

2017-10-01

In May of 2011, NASA selected the Origins, Spectral Interpretation, Resource Identification, and Security- Regolith Explorer (OSIRIS-REx) asteroid sample return mission as the third mission in the New Frontiers program. The other two New Frontiers missions are New Horizons, which explored Pluto during a flyby in July 2015 and is on its way for a flyby of Kuiper Belt object 2014 MU69 on January 1, 2019, and Juno, an orbiting mission that is studying the origin, evolution, and internal structure of Jupiter. The spacecraft departed for near-Earth asteroid (101955) Bennu aboard an United Launch Alliance Atlas V 411 evolved expendable launch vehicle at 7:05 p.m. EDT on September 8, 2016, on a seven-year journey to return samples from Bennu. The spacecraft is on an outbound-cruise trajectory that will result in a rendezvous with Bennu in November 2018. The science instruments on the spacecraft will survey Bennu to measure its physical, geological, and chemical properties, and the team will use these data to select a site on the surface to collect at least 60 g of asteroid regolith. The team will also analyze the remote-sensing data to perform a detailed study of the sample site for context, assess Bennu's resource potential, refine estimates of its impact probability with Earth, and provide ground-truth data for the extensive astronomical data set collected on this asteroid. The spacecraft will leave Bennu in 2021 and return the sample to the Utah Test and Training Range (UTTR) on September 24, 2023.

Integrated Modeling of Spacecraft Touch-and-Go Sampling

NASA Technical Reports Server (NTRS)

Quadrelli, Marco

2009-01-01

An integrated modeling tool has been developed to include multi-body dynamics, orbital dynamics, and touch-and-go dynamics for spacecraft covering three types of end-effectors: a sticky pad, a brush-wheel sampler, and a pellet gun. Several multi-body models of a free-flying spacecraft with a multi-link manipulator driving these end-effectors have been tested with typical contact conditions arising when the manipulator arm is to sample the surface of an asteroidal body. The test data have been infused directly into the dynamics formulation including such information as the mass collected as a function of end-effector longitudinal speed for the brush-wheel and sticky-pad samplers, and the mass collected as a function of projectile speed for the pellet gun sampler. These data represent the realistic behavior of the end effector while in contact with a surface, and represent a low-order model of more complex contact conditions that otherwise would have to be simulated. Numerical results demonstrate the adequacy of these multibody models for spacecraft and manipulator- arm control design. The work contributes to the development of a touch-and-go testbed for small body exploration, denoted as the GREX Testbed (GN&C for Rendezvous-based EXploration). The GREX testbed addresses the key issues involved in landing on an asteroidal body or comet; namely, a complex, low-gravity field; partially known terrain properties; possible comet outgassing; dust ejection; and navigating to a safe and scientifically desirable zone.

A manned maneuvering unit proximity operations planning and flight guidance display and control system

NASA Technical Reports Server (NTRS)

Gershzohn, Gary R.; Sirko, Robert J.; Zimmerman, K.; Jones, A. D.

1990-01-01

This task concerns the design, development, testing, and evaluation of a new proximity operations planning and flight guidance display and control system for manned space operations. A forecast, derivative manned maneuvering unit (MMU) was identified as a candidate for the application of a color, highway-in-the-sky display format for the presentation of flight guidance information. A silicon graphics 4D/20-based simulation is being developed to design and test display formats and operations concepts. The simulation includes the following: (1) real-time color graphics generation to provide realistic, dynamic flight guidance displays and control characteristics; (2) real-time graphics generation of spacecraft trajectories; (3) MMU flight dynamics and control characteristics; (4) control algorithms for rotational and translational hand controllers; (5) orbital mechanics effects for rendezvous and chase spacecraft; (6) inclusion of appropriate navigation aids; and (7) measurement of subject performance. The flight planning system under development provides for: (1) selection of appropriate operational modes, including minimum cost, optimum cost, minimum time, and specified ETA; (2) automatic calculation of rendezvous trajectories, en route times, and fuel requirements; (3) and provisions for manual override. Man/machine function allocations in planning and en route flight segments are being evaluated. Planning and en route data are presented on one screen composed of two windows: (1) a map display presenting a view perpendicular to the orbital plane, depicting flight planning trajectory and time data attitude display presenting attitude and course data for use en route; and (2) an attitude display presenting local vertical-local horizontal attitude data superimposed on a highway-in-the-sky or flight channel representation of the flight planned course. Both display formats are presented while the MMU is en route. In addition to these displays, several original display elements are being developed, including a 3DOF flight detector for attitude commanding, a different flight detector for translation commands, and a pictorial representation of velocity deviations.

Innovative Strategies for Asteroid Precursor Exploration

NASA Astrophysics Data System (ADS)

Klaus, K.; Lawrence, S.; Elsperman, M. S.; Smith, D. B.

2011-12-01

Introduction: Our ambitions for space exploration have outpaced our ability to afford frequent visits to targets of interest. Launch costs and development times continue to increase for getting large space craft to deep space. This particularly affects workforce development and imperils opportunities for new development starts. The time has come to leverage technology advances (including advances in autonomous operation and propulsion technology) to reduce the cost and increase the flight rate of planetary missions, while actively developing a scientific and engineering workforce to achieve national space objectives. Background: As demonstrated by the 1994 Clementine mission, planetary exploration missions maximizing off-the-shelf components to obtain a focused set of measurement objectives can make meaningful contributions to advancing the frontiers of space exploration by achieving numerous science and exploration objectives. Near Earth Objects [NEOs] are interesting candidates for missions of this nature. While results from recent missions (i.e., Hayabusa, NEAR, Dawn) have dramatically increased our understanding of asteroids, important questions remain. For example, characterizing the properties of asteroid regolith is an important consideration for understanding telescopic observations of asteroids, as well as preparing for future asteroid human exploration. Spacecraft Concepts: There are many candidate target asteroids that are attainable with our concept. We envision a "mothership" carrying 2-3 nanosats to the target. The nanosats would serve as in-situ explorers. The spacecraft is notionally designed for launch on a Taurus II. Our study intends on validating the concept and our notional spacecraft design will be refined and presented. The current dry mass with nanosats is estimated to be 750kg. The 1999 JU3 mission concept is a rendezvous with a 950 kg of initial spacecraft mass, launched to a C3 of 4 km2/s2. Subtracting the spacecraft dry mass from the initial mass gives a propellant loading of 200 kg. The solution for this case required 115.3 kg of propellant, leaving a 42% propellant margin. Science Instrumentation: Key objectives of this notional asteroid explorer would include: (1) high-resolution surface topography; (2) characterization surface composition and mineralogy; (3) quantification of the radiation environment near an NEO; and (4) mechanical properties of surface, if a touchdown takes place. Each nanosat would notionally contain a stereo camera for navigation, an alpha proton x-ray spectrometer to make measurements of the surface chemistry, and a microscopic imaging system to characterize the particle size distribution of asteroid regolith; multiple nanosats would provided redundancy for the in-situ surface characterization phase of the mission and enable a rudimentary gravity map through radio signal tracking.

Rendezvous and Docking for Space Exploration

NASA Technical Reports Server (NTRS)

Machula, M. F.; Crain, T.; Sandhoo, G. S.

2005-01-01

To achieve the exploration goals, new approaches to exploration are being envisioned that include robotic networks, modular systems, pre-positioned propellants and in-space assembly in Earth orbit, Lunar orbit and other locations around the cosmos. A fundamental requirement for rendezvous and docking to accomplish in-space assembly exists in each of these locations. While existing systems and technologies can accomplish rendezvous and docking in low earth orbit, and rendezvous and docking with crewed systems has been successfully accomplished in low lunar orbit, our capability must extend toward autonomous rendezvous and docking. To meet the needs of the exploration vision in-space assembly requiring both crewed and uncrewed vehicles will be an integral part of the exploration architecture. This paper focuses on the intelligent application of autonomous rendezvous and docking technologies to meet the needs of that architecture. It also describes key technology investments that will increase the exploration program's ability to ensure mission success, regardless of whether the rendezvous are fully automated or have humans in the loop.

Hyperbolic Rendezvous at Mars: Risk Assessments and Mitigation Strategies

NASA Technical Reports Server (NTRS)

Jedrey, Ricky; Landau, Damon; Whitley, Ryan

2015-01-01

Given the current interest in the use of flyby trajectories for human Mars exploration, a key requirement is the capability to execute hyperbolic rendezvous. Hyperbolic rendezvous is used to transport crew from a Mars centered orbit, to a transiting Earth bound habitat that does a flyby. Representative cases are taken from future potential missions of this type, and a thorough sensitivity analysis of the hyperbolic rendezvous phase is performed. This includes early engine cutoff, missed burn times, and burn misalignment. A finite burn engine model is applied that assumes the hyperbolic rendezvous phase is done with at least two burns.

Cislunar Near Rectilinear Halo Orbit for Human Space Exploration

NASA Technical Reports Server (NTRS)

Whitley, Ryan; Martinez, Roland; Condon, Gerald; Williams, Jacob; Lee, David; Davis, Diane; Barton, Gregg; Bhatt, Sagar; Jang, Jiann-Woei; Clark, Fred;

Dragon Spacecraft during Flyunder Operations

NASA Image and Video Library

2012-05-24

ISS031-E-067404 (24 May 2012) --- The SpaceX Dragon commercial cargo craft approaches the International Space Station on May 24, 2012 for a series of tests to clear it for its final rendezvous and grapple on May 25. At 3:58 a.m. (EDT), Dragon performed a height adjust burn to bring it to a path 2.4 kilometers below the station. During this “fly-under,” Dragon established UHF communication with the station using its Commercial Orbital Transportation Services (COTS) Ultra-high frequency Communication Unit (CUCU). Dragon performed a test of its Relative GPS system, which uses the relative positions of the spacecraft to the space station to determine its location. On May 25, Expedition 31 Flight Engineers Don Pettit and Andre Kuipers will use the Canadarm2 robotic arm to grapple the supply ship about 8:06 a.m., with the berthing to the Earth-facing side of the station’s Harmony node following about 11:20 a.m. Dragon is scheduled to spend about a week docked with the station before returning to Earth on May 31 for retrieval.

Dragon Spacecraft during Flyunder Operations

NASA Image and Video Library

2012-05-24

ISS031-E-067345 (24 May 2012) --- The SpaceX Dragon commercial cargo craft approaches the International Space Station on May 24, 2012 for a series of tests to clear it for its final rendezvous and grapple on May 25. At 3:58 a.m. (EDT), Dragon performed a height adjust burn to bring it to a path 2.4 kilometers below the station. During this “fly-under,” Dragon established UHF communication with the station using its Commercial Orbital Transportation Services (COTS) Ultra-high frequency Communication Unit (CUCU). Dragon performed a test of its Relative GPS system, which uses the relative positions of the spacecraft to the space station to determine its location. On May 25, Expedition 31 Flight Engineers Don Pettit and Andre Kuipers will use the Canadarm2 robotic arm to grapple the supply ship about 8:06 a.m., with the berthing to the Earth-facing side of the station’s Harmony node following about 11:20 a.m. Dragon is scheduled to spend about a week docked with the station before returning to Earth on May 31 for retrieval.

Dragon Spacecraft during Flyunder Operations

NASA Image and Video Library

2012-05-24

ISS031-E-067328 (24 May 2012) --- The SpaceX Dragon commercial cargo craft approaches the International Space Station on May 24, 2012 for a series of tests to clear it for its final rendezvous and grapple on May 25. At 3:58 a.m. (EDT), Dragon performed a height adjust burn to bring it to a path 2.4 kilometers below the station. During this “fly-under,” Dragon established UHF communication with the station using its Commercial Orbital Transportation Services (COTS) Ultra-high frequency Communication Unit (CUCU). Dragon performed a test of its Relative GPS system, which uses the relative positions of the spacecraft to the space station to determine its location. On May 25, Expedition 31 Flight Engineers Don Pettit and Andre Kuipers will use the Canadarm2 robotic arm to grapple the supply ship about 8:06 a.m., with the berthing to the Earth-facing side of the station’s Harmony node following about 11:20 a.m. Dragon is scheduled to spend about a week docked with the station before returning to Earth on May 31 for retrieval.

Dragon Spacecraft during Flyunder Operations

NASA Image and Video Library

2012-05-24

ISS031-E-067401 (24 May 2012) --- The SpaceX Dragon commercial cargo craft approaches the International Space Station on May 24, 2012 for a series of tests to clear it for its final rendezvous and grapple on May 25. At 3:58 a.m. (EDT), Dragon performed a height adjust burn to bring it to a path 2.4 kilometers below the station. During this “fly-under,” Dragon established UHF communication with the station using its Commercial Orbital Transportation Services (COTS) Ultra-high frequency Communication Unit (CUCU). Dragon performed a test of its Relative GPS system, which uses the relative positions of the spacecraft to the space station to determine its location. On May 25, Expedition 31 Flight Engineers Don Pettit and Andre Kuipers will use the Canadarm2 robotic arm to grapple the supply ship about 8:06 a.m., with the berthing to the Earth-facing side of the station’s Harmony node following about 11:20 a.m. Dragon is scheduled to spend about a week docked with the station before returning to Earth on May 31 for retrieval.

Highest Resolution Topography of 433 Eros and Implications for MUSES-C

NASA Technical Reports Server (NTRS)

Cheng, A. F.; Barnouin-Jha, O.

2003-01-01

The highest resolution observations of surface morphology and topography at asteroid 433 Eros were obtained by the Near Earth Asteroid Rendezvous (NEAR) Shoemaker spacecraft on 12 February 2001, as it landed within a ponded deposit on Eros. Coordinated observations were obtained by the imager and the laser rangefinder, at best image resolution of 1 cm/pixel and best topographic resolution of 0.4 m. The NEAR landing datasets provide unique information on rock size and height distributions and regolith processes. Rocks and soil can be distinguished photometrically, suggesting that bare rock is indeed exposed. The NEAR landing data are the only data at sufficient resolution to be relevant to hazard assessment on future landed missions to asteroids, such as the MUSES-C mission which will land on asteroid 25143 (1998 SF36) in order to obtain samples. In a typical region just outside the pond where NEAR landed, the areal coverage by resolved positive topographic features is 18%. At least one topographic feature in the vicinity of the NEAR landing site would have been hazardous for a spacecraft.

Update on PISCES

NASA Technical Reports Server (NTRS)

Pearson, Don; Hamm, Dustin; Kubena, Brian; Weaver, Jonathan K.

2010-01-01

An updated version of the Platform Independent Software Components for the Exploration of Space (PISCES) software library is available. A previous version was reported in Library for Developing Spacecraft-Mission-Planning Software (MSC-22983), NASA Tech Briefs, Vol. 25, No. 7 (July 2001), page 52. To recapitulate: This software provides for Web-based, collaborative development of computer programs for planning trajectories and trajectory- related aspects of spacecraft-mission design. The library was built using state-of-the-art object-oriented concepts and software-development methodologies. The components of PISCES include Java-language application programs arranged in a hierarchy of classes that facilitates the reuse of the components. As its full name suggests, the PISCES library affords platform-independence: The Java language makes it possible to use the classes and application programs with a Java virtual machine, which is available in most Web-browser programs. Another advantage is expandability: Object orientation facilitates expansion of the library through creation of a new class. Improvements in the library since the previous version include development of orbital-maneuver- planning and rendezvous-launch-window application programs, enhancement of capabilities for propagation of orbits, and development of a desktop user interface.

The NASA Engineering and Safety Center (NESC) GN and C Technical Discipline Team (TDT): Its Purpose, Practices and Experiences

NASA Technical Reports Server (NTRS)

Dennehy, Cornelius J.

2008-01-01

This paper will briefly define the vision, mission, and purpose of the NESC organization. The role of the GN&C TDT will then be described in detail along with an overview of how this team operates and engages in its objective engineering and safety assessments of critical NASA projects. This paper will then describe key issues and findings from several of the recent GN&C-related independent assessments and consultations performed and/or supported by the NESC GN&C TDT. Among the examples of the GN&C TDT s work that will be addressed in this paper are the following: the Space Shuttle Orbiter Repair Maneuver (ORM) assessment, the ISS CMG failure root cause assessment, the Demonstration of Autonomous Rendezvous Technologies (DART) spacecraft mishap consultation, the Phoenix Mars lander thruster-based controllability consultation, the NASA in-house Crew Exploration Vehicle (CEV) Smart Buyer assessment and the assessment of key engineering considerations for the Design, Development, Test & Evaluation (DDT&E) of robust and reliable GN&C systems for human-rated spacecraft.

An automated rendezvous and capture system design concept for the cargo transfer vehicle and Space Station Freedom

NASA Technical Reports Server (NTRS)

Fuchs, Ron; Marsh, Steven

1991-01-01

A rendezvous sensor system concept was developed for the cargo transfer vehicle (CTV) to autonomously rendezvous with and be captured by Space Station Freedom (SSF). The development of requirements, the design of a unique Lockheed developed sensor concept to meet these requirements, and the system design to place this sensor on the CTV and rendezvous with the SSF are described .

Investigating Trojan Asteroids at the L4/L5 Sun-Earth Lagrange Points

NASA Technical Reports Server (NTRS)

John, K. K.; Graham, L. D.; Abell, P. A.

2015-01-01

Investigations of Earth's Trojan asteroids will have benefits for science, exploration, and resource utilization. By sending a small spacecraft to the Sun-Earth L4 or L5 Lagrange points to investigate near-Earth objects, Earth's Trojan population can be better understood. This could lead to future missions for larger precursor spacecraft as well as human missions. The presence of objects in the Sun-Earth L4 and L5 Lagrange points has long been suspected, and in 2010 NASA's Wide-field Infrared Survey Explorer (WISE) detected a 300 m object. To investigate these Earth Trojan asteroid objects, it is both essential and feasible to send spacecraft to these regions. By exploring a wide field area, a small spacecraft equipped with an IR camera could hunt for Trojan asteroids and other Earth co-orbiting objects at the L4 or L5 Lagrange points in the near-term. By surveying the region, a zeroth-order approximation of the number of objects could be obtained with some rough constraints on their diameters, which may lead to the identification of potential candidates for further study. This would serve as a precursor for additional future robotic and human exploration targets. Depending on the inclination of these potential objects, they could be used as proving areas for future missions in the sense that the delta-V's to get to these targets are relatively low as compared to other rendezvous missions. They can serve as platforms for extended operations in deep space while interacting with a natural object in microgravity. Theoretically, such low inclination Earth Trojan asteroids exist. By sending a spacecraft to L4 or L5, these likely and potentially accessible targets could be identified.

Observations and Characterization of Binary Near-Earth Asteroid 65803 Didymos, the Target of the AIDA Mission

NASA Astrophysics Data System (ADS)

Naidu, S.; Benner, L.; Brozovic, M.; Ostro, S. J.; Nolan, M. C.; Margot, J. L.; Giorgini, J. D.; Magri, C.; Pravec, P.; Scheirich, P.; Scheeres, D. J.; Hirabayashi, M.

2016-12-01

Binary near-Earth asteroid 65803 Didymos is the target of the proposed Asteroid Impact and Deflection Assessment (AIDA) space mission. The mission consists of two spacecraft, the Demonstration for Autonomous Rendezvous Technology (DART) spacecraft that will impact the asteroid's satellite and the Asteroid Impact Mission (AIM) spacecraft that will observe the impact. We used radar observations obtained at Arecibo and Goldstone in 2003, and lightcurve data from Pravec et al. (2006) to model the shapes, sizes, and spin states of the components. The primary is top shaped and has an equatorial ridge similar to the one seen on 2000 DP107 (Naidu et al. 2015). A 300 m long flat region is also seen along the equator. The primary has an equivalent diameter of 780 m (+/- 10 %) and its extents along the principal axes are 826 m, 813 m, and 786 m (10% uncertainties). It has a spin period of 2.2600 +/- 0.0001 h. A grid search for the spin pole resulted in the best fit at ecliptic (longitude, latitude) = (296, +71) degrees (+/- 15 degrees). This estimate is consistent with the spin pole being aligned to the binary orbit normal at (310, -84) degrees. Dividing the primary mass of 5.24e11 kg (Fang & Margot 2012) by the model volume we estimate a bulk density of 2100 kg m-3 (+/- 30 %). We summed multiple radar runs to estimate the range and Doppler extents of the satellite. We estimated the motion in successive images and used a shift-and-sum technique to mitigate smearing due to translational motion. This boosted the SNRs and allowed us to obtain size and bandwidth estimates of the satellite. The visible range extent of the satellite is roughly 60-75 m at the 15 m resolution of the Arecibo images. Assuming that the true extent is twice the visible extent, we obtain a diameter estimate of 120-150 m. The bandwidth of the satellite suggests a spin period between 9-12 h that is consistent with the orbit period of 11.9 hours and with synchronous rotation.

Mission commander James Wetherbee on the forward flight deck

NASA Image and Video Library

1995-02-03

STS063-06-027 (3-11 Feb 1995) --- Seated at the commander's station on the Space Shuttle Discovery's flight deck, astronaut James D. Wetherbee, commander, was photographed by a crew mate during early phases of the STS-63 mission. A great deal of time was spent during the first few days of the mission to check a leaky thruster, which could have had a negative influence on rendezvous operations with Russia's Mir Space Station. As it turned out, all the related problems were solved and the two spacecraft succeded in achieving close proximity operations. Others onboard the Discovery were astronauts Eileen M. Collins, pilot; Bernard A. Harris Jr., payload commander; and mission specialists C. Michael Foale, Janice E. Voss, and Russian cosmonaut Vladimir G. Titov.

The Asteroid Impact Mission - Deflection Demonstration (AIM - D2)

NASA Astrophysics Data System (ADS)

Küppers, M.; Michel, P.; Carnelli, I.

2017-09-01

The Asteroid Impact Mission (AIM) is ESA's contribution to the international Asteroid Impact Deflection Assessment (AIDA) cooperation, targeting the demonstration of deflection of a hazardous near-earth asteroid. AIM will also be the first in-depth investigation of a binary asteroid and make measurements that are relevant for the preparation of asteroid resource utilisation. AIM is foreseen to rendezvous with the binary near-Earth asteroid (65803) Didymos and to observe the system before, during, and after the impact of NASA's Double Asteroid Redirection Test (DART) spacecraft. Here we describe the observations to be done by the simplified version Asteroid Impact Mission - Deflection Demonstration (AIM-D2) and show that most of the original AIM objectives can still be achieved.

The NEAR Multispectral Imager.

NASA Astrophysics Data System (ADS)

Hawkins, S. E., III

1998-06-01

Multispectral Imager, one of the primary instruments on the Near Earth Asteroid Rendezvous (NEAR) spacecraft, uses a five-element refractive optics telescope, an eight-position filter wheel, and a charge-coupled device detector to acquire images over its sensitive wavelength range of ≍400 - 1100 nm. The primary science objectives of the Multispectral Imager are to determine the morphology and composition of the surface of asteroid 433 Eros. The camera will have a critical role in navigating to the asteroid. Seven narrowband spectral filters have been selected to provide multicolor imaging for comparative studies with previous observations of asteroids in the same class as Eros. The eighth filter is broadband and will be used for optical navigation. An overview of the instrument is presented, and design parameters and tradeoffs are discussed.

Selected tether applications in space: Phase 2. Executive summary

NASA Technical Reports Server (NTRS)

Thorson, M. H.; Lippy, L. J.

1985-01-01

The application of tether technology has the potential to increase the overall performance efficiency and capability of the integrated space operations and transportation systems through the decade of the 90s. The primary concepts for which significant economic benefits were identified are dependent on the space station as a storage device for angular momentum and as an operating base for the tether system. Concepts examined include: (1) tether deorbit of shuttle from space station; (2) tethered orbit insertion of a spacecraft from shuttle; (3) tethered platform deployed from space station; (4) tether-effected rendezvous of an OMV with a returning OTV; (5) electrodynamic tether as an auxiliary power source for space station; and (6) tether assisted launch of an OTV mission from space station.

Users manual for the IMA program

NASA Technical Reports Server (NTRS)

Williams, D. F.

1991-01-01

The Impulsive Mission Analysis (IMA) computer program provides a user-friendly means of designing a complete Earth-orbital mission profile using an 80386-based microcomputer. The IMA program produces a trajectory summary, an output file for use by the new Simplex Computation of Optimum Orbital Trajectories (SCOOT) program, and several graphics, including ground tracks on a world map, altitude profiles, relative motion plots, and sunlight/communication timelines. The user can design missions using any combination of three basic types of mission segments: double co-eliptic rendezvous, payload delivery, and payload de-orbit/spacecraft recovery. Each mission segment is divided into one or more transfers, and each transfer is divided into one or more legs, each leg consisting of a coast arc followed by a burn arc.

Flight Dynamics and GN&C for Spacecraft Servicing Missions

NASA Technical Reports Server (NTRS)

Naasz, Bo; Zimpfer, Doug; Barrington, Ray; Mulder, Tom

2010-01-01

Future human exploration missions and commercial opportunities will be enabled through In-space assembly and satellite servicing. Several recent efforts have developed technologies and capabilities to support these exciting future missions, including advances in flight dynamics and Guidance, Navigation and Control. The Space Shuttle has demonstrated significant capabilities for crewed servicing of the Hubble Space Telescope (HST) and assembly of the International Space Station (ISS). Following the Columbia disaster NASA made significant progress in developing a robotic mission to service the HST. The DARPA Orbital Express mission demonstrated automated rendezvous and capture, In-space propellant transfer, and commodity replacement. This paper will provide a summary of the recent technology developments and lessons learned, and provide a focus for potential future missions.

PROTOCOL - APOLLO-SOYUZ TEST PROJECT (ASTP) - TOUR - WASHINGTON, DC

NASA Image and Video Library

1974-09-07

S74-29892 (7 Sept. 1974) --- President Gerald R. Ford removes the Soviet Soyuz spacecraft model from a model set depicting the 1975 Apollo-Soyuz Test Project, an Earth orbital docking and rendezvous mission involving crewmen from the U.S. and USSR, who visited Mr. Ford at the White House. The cosmonauts and astronauts are, left to right, Vladimir A. Shatalov, Chief, Cosmonaut Training; Valeriy N. Kubasov, ASTP Soviet engineer; Aleksey A. Leonov, ASTP Soviet crew commander; Thomas P. Stafford, ASTP American crew commander; Donald K. Slayton, American crew?s docking module pilot; and Vance D. Brand, command module pilot for the U.S. team. Dr. George M. Low, Deputy Administrator, National Aeronautics and Space Administration, is partially obscured behind Mr. Ford.

Short-range/Long-range Integrated Target (SLIT) for Video Guidance Sensor Rendezvous and Docking

NASA Technical Reports Server (NTRS)

Roe, Fred D. (Inventor); Bryan, Thomas C. (Inventor)

2009-01-01

A laser target reflector assembly for mounting upon spacecraft having a long-range reflector array formed from a plurality of unfiltered light reflectors embedded in an array pattern upon a hemispherical reflector disposed upon a mounting plate. The reflector assembly also includes a short-range reflector array positioned upon the mounting body proximate to the long-range reflector array. The short-range reflector array includes three filtered light reflectors positioned upon extensions from the mounting body. The three filtered light reflectors retro-reflect substantially all incident light rays that are transmissive by their monochromatic filters and received by the three filtered light reflectors. In one embodiment the short-range reflector array is embedded within the hemispherical reflector,

Development of Tools and Techniques for Processing STORRM Flight Data

NASA Technical Reports Server (NTRS)

Robinson, Shane; D'Souza, Christopher

2011-01-01

While at JSC for the summer of 2011, I was assigned to work on the sensor test for Orion relative-navigation risk mitigation (STORRM) development test objective (DTO). The STORRM DTO was flown on-board Endeavor during STS-134. The objective of the STORRM DTO is to test the visual navigation system (VNS), which will be used as the primary relative navigation sensor for the Orion spacecraft. The VNS is a flash lidar system intended to provide both line of sight and range information during rendezvous and proximity operations. The STORRM DTO also serves as a testbed for the high-resolution docking camera. This docking camera will be used to provide piloting cues for the crew during proximity operations. These instruments were mounted next to the trajectory control sensor (TCS) in Endeavour s payload bay. My principle objective for the summer was to generate a best estimated trajectory (BET) for Endeavor using the flight data collected by the VNS during rendezvous and the unprecedented re-rendezvous with the ISS. I processed the raw images from the VNS to produce range and bearing measurements. I then aggregated these measurements and extracted the measurements corresponding to individual reflectors. I combined the information contained in these measurements with data from the Endeavour's inertial sensors using Kalman smoothing techniques to ultimately produce a BET. This work culminated with a final presentation of the result to division management. Development of this tool required that traditional linear smoothing techniques be modified in a novel fashion to permit for the inclusion of non-linear measurements. This internship has greatly helped me further my career by providing exposure to real engineering projects. I also have benefited immensely from the mentorship of the engineers working on these projects. Many of the lessons I learned and experiences I had are of particular value because then can only be found in a place like JSC.

Method for deploying multiple spacecraft

NASA Technical Reports Server (NTRS)

Sharer, Peter J. (Inventor)

2007-01-01

A method for deploying multiple spacecraft is disclosed. The method can be used in a situation where a first celestial body is being orbited by a second celestial body. The spacecraft are loaded onto a single spaceship that contains the multiple spacecraft and the spacecraft is launched from the second celestial body towards a third celestial body. The spacecraft are separated from each other while in route to the third celestial body. Each of the spacecraft is then subjected to the gravitational field of the third celestial body and each of the spacecraft assumes a different, independent orbit about the first celestial body. In those situations where the spacecraft are launched from Earth, the Sun can act as the first celestial body, the Earth can act as the second celestial body and the Moon can act as the third celestial body.

Omnidirectional angle constraint based dynamic six-degree-of-freedom measurement for spacecraft rendezvous and docking simulation

NASA Astrophysics Data System (ADS)

Shi, Shendong; Yang, Linghui; Lin, Jiarui; Ren, Yongjie; Guo, Siyang; Zhu, Jigui

2018-04-01

In this paper we present a novel omnidirectional angle constraint based method for dynamic 6-DOF (six-degree-of-freedom) measurement. A photoelectric scanning measurement network is employed whose photoelectric receivers are fixed on the measured target. They are in a loop distribution and receive signals from rotating transmitters. Each receiver indicates an angle constraint direction. Therefore, omnidirectional angle constraints can be constructed in each rotation cycle. By solving the constrained optimization problem, 6-DOF information can be obtained, which is independent of traditional rigid coordinate system transformation. For the dynamic error caused by the measurement principle, we present an interpolation method for error reduction. Accuracy testing is performed in an 8  ×  8 m measurement area with four transmitters. The experimental results show that the dynamic orientation RMSEs (root-mean-square errors) are reduced from 0.077° to 0.044°, 0.040° to 0.030° and 0.032° to 0.015° in the X, Y, and Z axes, respectively. The dynamic position RMSE is reduced from 0.65 mm to 0.24 mm. This method is applied during the final approach phase in the rendezvous and docking simulation. Experiments under different conditions are performed in a 40  ×  30 m area, and the method is verified to be effective.

The GNC Measurement System for the Automated Transfer Vehicle

NASA Astrophysics Data System (ADS)

Roux, Y.; da Cunha, P.

The Automated Transfer Vehicle (ATV) is a European Space Agency (ESA) funded spacecraft developed by EADS Space Transportation as prime contractor for the space segment together with major European industrial partners, in the frame of the International Space Station (ISS). Its mission objective is threefold : to supply the station with fret and propellant, to reboost ISS to a higher orbit and to dispose of waste from the station. The ATV first flight, called Jules Verne and planned on 2005, will be the first European Vehicle to perform an orbital rendezvous. The GNC Measurement System (GMS) is the ATV on board function in charge of the measurement data collection and preconditioning for the navigation, guidance and control (GNC) algorithms. The GMS is made up of hardware which are the navigation sensors (with a certain level of hardware redundancy for each of them), and of an on-board software that manages, monitors and performs consistency checks to detect and isolate potential sensor failures. The GMS relies on six kinds of navigation sensors, used during various phases of the mission : the gyrometers assembly (GYRA), the accelerometers assembly (ACCA), the star trackers (STR), the GPS receivers, the telegoniometers (TGM) and the videometers (VDM), the last two being used for the final rendezvous phase. The GMS function is developed by EADS Space Transportation together with other industrial partners: EADS Astrium, EADS Sodern, Laben and Dasa Jena Optronik.

General Methodology for Designing Spacecraft Trajectories

NASA Technical Reports Server (NTRS)

Condon, Gerald; Ocampo, Cesar; Mathur, Ravishankar; Morcos, Fady; Senent, Juan; Williams, Jacob; Davis, Elizabeth C.

2012-01-01

A methodology for designing spacecraft trajectories in any gravitational environment within the solar system has been developed. The methodology facilitates modeling and optimization for problems ranging from that of a single spacecraft orbiting a single celestial body to that of a mission involving multiple spacecraft and multiple propulsion systems operating in gravitational fields of multiple celestial bodies. The methodology consolidates almost all spacecraft trajectory design and optimization problems into a single conceptual framework requiring solution of either a system of nonlinear equations or a parameter-optimization problem with equality and/or inequality constraints.

Rosetta Lander - Philae: preparations for landing on comet 67P/Churyumov-Gerasimenko

NASA Astrophysics Data System (ADS)

Ulamec, S.; Biele, J.; Jurado, E.; Gaudon, P.; Geurts, K.

2013-12-01

Rosetta is a Cornerstone Mission of the ESA Horizon 2000 programme. It is going to rendezvous with comet 67P/Churyumov-Gerasimenko after a ten year cruise and will study both its nucleus and coma with an orbiting spacecraft as well as with a Lander, Philae, that has been designed to land softly on the comet nucleus. Aboard Philae, a payload consisting of ten scientific instruments will perform in-situ studies of the cometary material. Philae will be separated from the mother spacecraft from a dedicated delivery trajectory. It then descends, ballistically, to the surface of the comet, stabilized with an internal flywheel. At touch-down anchoring harpoons will be fired and a damping mechanism within the landing gear will provide the lander from re-bouncing. Currently the characteristics of the nucleus of the comet are hardly known. Mapping with the orbiter cameras (shape, slopes, surface roughness) and essential measurements like gravity field, state of rotation or outgassing parameters can only be performed after arrival of the main spacecraft, between May and October 2014. These data will be used for selecting a landing site and defining the detailed landing strategy. Landing is foreseen for November 2014 at a heliocentric distance of 3 AU. The paper describes the Rosetta Lander system and its payload, but emphasizes on the preparations for landing, the landing site selection process and the planned operational timeline.

Analysis of laser radar measurements of the asteroid 433 Eros

NASA Astrophysics Data System (ADS)

Cole, Timothy D.; Zuber, Maria T.; Neuman, Greg; Cheng, Andrew F.; Reiter, R. Alan; Guo, Yanping; Smith, David E.

2001-09-01

After a 5-year mission, a 4-year transit followed by a one-year mission orbiting the asteroid 433 Eros, the Near-Earth Asteroid Rendezvous-Shoemaker (NEAR) spacecraft made a controlled landing onto the asteroid's surface on 12 February 2001. Onboard the spacecraft, the NEAR Laser Rangefinder (NLR) facility instrument had gathered over 11 million measurements, providing a spatially dense, high-resolution, topographical map of Eros. This topographic data, combined with Doppler tracking data for the spacecraft, enabled the determination of the asteroid's shape, mass, and density thereby contributing to understanding the internal structure and collisional evolution of Eros. NLR data indicate that Eros is a consolidated body with a complex shape dominated by collisions. The offset between the asteroid's center of mass and center of figure indicates a small deviation from a homogeneous internal structure that is most simply explained by variations in mechanical structure. Regional-scale relief and slope distributions show evidence for control of some topography by a competent substrate. It was found that pulse dilation was the major source of uncertainty in single-shot range measurements from the NLR, and that this uncertainty remains consistent with the overall 6-m range measurement system accuracy for NEAR. Analysis of NLR data fully quantified the geodynamic nature of this planetesimal, ergo, illustrating the utility of laser altimetry for remote sensing.

The opearation and scientific data of MINERVA rover in Hayabusa mission

NASA Astrophysics Data System (ADS)

Yoshimitsu, T.; Kubota, T.; Nakatani, I.

section Introduction The asteroid explorer Hayabusa operated by Institute of Space and Astronautical Science ISAS JAXA made a rendezvous with the target asteroid Itokawa in September 2005 after more than two years interplanetary cruise since the launch in 2003 The spacecraft precisely observed the target from the vicinity of the asteroid for two months and then tried to land on the surface in order to get some fragments which will be brought back to the Earth The authors installed a experimental small rover named MINERVA into the spacecraft It was supposed to make a surface exploration after having been deployed onto the surface MINERVA is the smallest spacecraft with a weight of 591 g and a dimension of about 10 cm in size In this small body it is equipped with a moving ability over the micro-gravity environment on the asteroid and a few scientific instruments cameras and thermometers to characterize the surface MINERVA was deployed in November 2005 but it could not reach at the asteroid because the deployment was not done at the good timing Thus it became a artificial solar satellite and the surface exploration was not conducted It survived at least for 18 hours after the deployment while the obtained data were transmitted to the Earth via the mother spacecraft This paper describes the operation and the possible science from the obtained data by MINERVA section Deployment MINERVA was deployed in November 12 2005 by sending a command from the Earth when the rehearsal operation of the touchdown was in

Omni-directional Particle Detector (ODPD) on Tiangong-2 Spacecraft

NASA Astrophysics Data System (ADS)

Guohong, S.; Zhang, S.; Yang, X.; Wang, C.

2017-12-01

Tiangong-2 spacecraft is the second space laboratory independently developed by china after Tiangong-1, which was launched on 15 September 2016. It is also the first real space laboratory in china, which will be used to further validate the space rendezvous and docking technology and to carry out a series of space tests. The spacecraft's orbit is 350km height and 42° inclination. The omni-directional particle detector (ODPD) on Tiangong-2 spacecraft is a new instrument developed by China. Its goal is the anisotropy and energy spectra of space particles on manned space flight orbit. The ODPD measures the energy spectra and pitch angle distributions of high energy electrons and protons. It consists of one electron spectrum telescope, one proton spectrum telescope and sixteen directional flux telescopes. The ODPD is designed to measure the protons spectrum from 2.5MeV to 150MeV, electrons spectrum from 0.2MeV to 1.5MeV, the flux of electrons energy >200keV and protons energy>1.5MeV on 2∏ space, also the ODPD has a small sensor to measure the LET spectrum from 1-100MeV/cm2sr. The primary advantage can give the particle's pitch angle distributions at any time because of the sixteen flux telescopes arrange form 0 to 180 degree. This is the first paper dealing with ODPD data, so we firstly spend some time describing the instrument, its theory of operation and its calibration. Then we give the preliminary detecting results.

Expert system isssues in automated, autonomous space vehicle rendezvous

NASA Technical Reports Server (NTRS)

Goodwin, Mary Ann; Bochsler, Daniel C.

1987-01-01

The problems involved in automated autonomous rendezvous are briefly reviewed, and the Rendezvous Expert (RENEX) expert system is discussed with reference to its goals, approach used, and knowledge structure and contents. RENEX has been developed to support streamlining operations for the Space Shuttle and Space Station program and to aid definition of mission requirements for the autonomous portions of rendezvous for the Mars Surface Sample Return and Comet Nucleus Sample return unmanned missions. The experience with REMEX to date and recommendations for further development are presented.

Remote Spacecraft Attitude Control by Coulomb Charging

NASA Astrophysics Data System (ADS)

Stevenson, Daan

The possibility of inter-spacecraft collisions is a serious concern at Geosynchronous altitudes, where many high-value assets operate in proximity to countless debris objects whose orbits experience no natural means of decay. The ability to rendezvous with these derelict satellites would enable active debris removal by servicing or repositioning missions, but docking procedures are generally inhibited by the large rotational momenta of uncontrolled satellites. Therefore, a contactless means of reducing the rotation rate of objects in the space environment is desired. This dissertation investigates the viability of Coulomb charging to achieve such remote spacecraft attitude control. If a servicing craft imposes absolute electric potentials on a nearby nonspherical debris object, it will impart electrostatic torques that can be used to gradually arrest the object's rotation. In order to simulate the relative motion of charged spacecraft with complex geometries, accurate but rapid knowledge of the Coulomb interactions is required. To this end, a new electrostatic force model called the Multi-Sphere Method (MSM) is developed. All aspects of the Coulomb de-spin concept are extensively analyzed and simulated using a system with simplified geometries and one dimensional rotation. First, appropriate control algorithms are developed to ensure that the nonlinear Coulomb torques arrest the rotation with guaranteed stability. Moreover, the complex interaction of the spacecraft with the plasma environment and charge control beams is modeled to determine what hardware requirements are necessary to achieve the desired electric potential levels. Lastly, the attitude dynamics and feedback control development is validated experimentally using a scaled down terrestrial testbed. High voltage power supplies control the potential on two nearby conductors, a stationary sphere and a freely rotating cylinder. The nonlinear feedback control algorithms developed above are implemented to achieve rotation rate and absolute attitude control. Collectively, these studies decisively validate the feasibility of Coulomb charging for remote spacecraft attitude control.

Automatic procedures generator for orbital rendezvous maneuver

NASA Technical Reports Server (NTRS)

Kohn, W.; Van Valkenburg, J. A.; Dunn, C. K.

1985-01-01

This paper describes the development of an expert system for defining and dynamically updating procedures for an orbital rendezvous maneuver. The product of the expert system is a procedure represented by a Moore automaton. The construction is recursive and driven by a simulation of the rendezvousing bodies.

Advanced missions safety. Volume 3: Appendices. Part 1: Space shuttle rescue capability

NASA Technical Reports Server (NTRS)

1972-01-01

The space shuttle rescue capability is analyzed as a part of the advanced mission safety study. The subjects discussed are: (1) mission evaluation, (2) shuttle configurations and performance, (3) performance of shuttle-launched tug system, (4) multiple pass grazing reentry from lunar orbit, (5) ground launched ascent and rendezvous time, (6) cost estimates, and (7) parallel-burn space shuttle configuration.

INSPACE CHEMICAL PROPULSION SYSTEMS AT NASA's MARSHALL SPACE FLIGHT CENTER: HERITAGE AND CAPABILITIES

NASA Technical Reports Server (NTRS)

McRight, P. S.; Sheehy, J. A.; Blevins, J. A.

2005-01-01

NASA s Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V rocket and the Space Shuttle external tank, solid rocket boosters, and main engines. This paper highlights a lesser known but very rich side of MSFC-its heritage in the development of in-space chemical propulsion systems and its current capabilities for spacecraft propulsion system development and chemical propulsion research. The historical narrative describes the flight development activities associated with upper stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology (DART), X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and multiple technology development activities. This paper also highlights MSFC s advanced chemical propulsion research capabilities, including an overview of the center s Propulsion Systems Department and ongoing activities. The authors highlight near-term and long-term technology challenges to which MSFC research and system development competencies are relevant. This paper concludes by assessing the value of the full range of aforementioned activities, strengths, and capabilities in light of NASA s exploration missions.

Robust adaptive relative position and attitude control for spacecraft autonomous proximity.

PubMed

Sun, Liang; Huo, Wei; Jiao, Zongxia

2016-07-01

This paper provides new results of the dynamical modeling and controller designing for autonomous close proximity phase during rendezvous and docking in the presence of kinematic couplings and model uncertainties. A globally defined relative motion mechanical model for close proximity operations is introduced firstly. Then, in spite of the kinematic couplings and thrust misalignment between relative rotation and relative translation, robust adaptive relative position and relative attitude controllers are designed successively. Finally, stability of the overall system is proved that the relative position and relative attitude are uniformly ultimately bounded, and the size of the ultimate bound can be regulated small enough by control system parameters. Performance of the controlled overall system is demonstrated via a representative numerical example. Copyright © 2016 ISA. Published by Elsevier Ltd. All rights reserved.

The Ion Propulsion System for the Asteroid Redirect Robotic Mission

NASA Technical Reports Server (NTRS)

Herman, Daniel A.; Santiago, Walter; Kamhawi, Hani; Polk, James E.; Snyder, John Steven; Hofer, Richard; Sekerak, Michael

2016-01-01

The Asteroid Redirect Robotic Mission is a Solar Electric Propulsion Technology Demonstration Mission (ARRM) whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of NASA's future beyond-low-Earth-orbit, human-crewed exploration plans. This presentation presents the conceptual design of the ARRM ion propulsion system, the status of the NASA in-house thruster and power processing development activities, the status of the planned technology maturation for the mission through flight hardware delivery, and the status of the mission formulation and spacecraft acquisition.