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NASA NTRS2021

FRESCO: A Framework for Spacecraft Systems Autonomy

Achieving the science exploration and defense goals of the following decades will require flight systems capable of operations with limited operator contact, system mode changes and retasking based on sensor data, and complex robotic operations. To support these capabilities, increasingly autonomous flight systems are required that can perform dedicated mission functions, e.g. payload targeting and communications, and system-level functions, e.g. planning and goal monitoring. Architecting an autonomous system requires a well-reasoned, self-consistent framework to avoid \textit{ad hoc} design choices that will introduce complexity and risk. The Framework for Robust Execution and Scheduling of Commands On-Board, FRESCO, is the result of lessons learned in developing a software architecture to enable autonomous solar system exploration. FRESCO generalizes this work to offer a modular, software-agnostic approach to developing verifiable architecture for autonomous space systems. FRESCO specifies guiding principles, functions, interfaces, and interactions from which mission-specific autonomous control architectures can be derived. FRESCO is a principled framework relying on explicit, state-based goal definitions, centralized management of state knowledge, clearly separated control boundaries, and hierarchical reasoning. Using components from FRESCO reference architecture, an autonomous decision-making architecture can be designed for spacecraft which can then be mapped to flight software architecture. FRESCO is flexibly defined to enable autonomous control of flight systems built using extensive software and hardware heritage. Finally, FRESCO-derived architectures support a spectrum of operator/spacecraft interactions, ranging from traditional commanding to goal-driven commanding with the ability to change mission goals autonomously. FRESCO has been used in defining the autonomy architectures for the ASTERIA mission and have been demonstrated in laboratory and software simulation for small body rendezvous and in-space servicing missions.

Kolcio, Ksenia↗

NASA NTRS2010

Liquid Acquisition Strategies for Exploration Missions: Current Status 2010

NASA is currently developing the propulsion system concepts for human exploration missions to the lunar surface. The propulsion concepts being investigated are considering the use of cryogenic propellants for the low gravity portion of the mission, that is, the lunar transit, lunar orbit insertion, lunar descent and the rendezvous in lunar orbit with a service module after ascent from the lunar surface. These propulsion concepts will require the vapor free delivery of the cryogenic propellants stored in the propulsion tanks to the exploration vehicles main propulsion system (MPS) engines and reaction control system (RCS) engines. Propellant management devices (PMD s) such as screen channel capillary liquid acquisition devices (LAD s), vanes and sponges currently are used for earth storable propellants in the Space Shuttle Orbiter OMS and RCS applications and spacecraft propulsion applications but only very limited propellant management capability exists for cryogenic propellants. NASA has begun a technology program to develop LAD cryogenic fluid management (CFM) technology through a government in-house ground test program of accurately measuring the bubble point delta-pressure for typical screen samples using LO2, LN2, LH2 and LCH4 as test fluids at various fluid temperatures and pressures. This presentation will document the CFM project s progress to date in concept designs, as well ground testing results.

Chato, David J.↗

NASA NTRS2007

Orion GN and C Overview and Architecture

The Crew Exploration Vehicle, named Orion, is a critical element in the Constellation Program to develop the transportation system needed to send humans back to the moon and then beyond. Lockheed Martin is the prime contractor for the Orion spacecraft, which is managed by the Johnson Space Center. The Orion GN&C sub-system is being jointly developed by NASA and Lockheed Martin through a mode team approach. The GN&C is a critical element of the Orion mission to carry astronauts to low earth orbit to service the International Space Station and then on later flights to transfer and return a crew of four to the moon. The Orion GN&C system must perform monitoring and abort functions during ascent, rendezvous and docking in both low earth and lunar orbits, perform uncrewed lunar loiter operations, perform trans earth injection and atmospheric entry and landing. The Orion also must be integrated with the Ares I Crew Launch Vehicle, the Earth Departure Stage of the Ares V and the Lunar Surface Access Module. This paper provides an overview of the Orion GN&C system. The functional capabilities of the Orion GN&C will be provided in the context of Constellation architecture, the key GN&C requirements will be summarized, the GN&C architecture will be presented, the development schedule and plans will summarized and finally conclusions will be presented.

Hu, Howard↗

NASA NTRS2008

Electric Propulsion System Selection Process for Interplanetary Missions

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.

Landau, Damon↗

NASA NTRS1972

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

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.

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NASA NTRS1982

Fuel optimal maneuvers of spacecraft about a circular orbit

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.

Carter, T. E.↗

NASA NTRS1984

Fuel-optimal maneuvers of a spacecraft relative to a point in circular orbit

Carter, T. E.↗

NASA NTRS1988

Outward to the Beginning: the CRAF and Cassini Missions of the Mariner Mark 2 Program

Two successive journeys will soon offer a perspective on the origin of the solar system and perhaps provide clues on the origin of life as well. The missions, the Comet Rendezvous Asteroid Flyby (CRAF) and Cassini (the Saturn orbiter/Titan probe), combine to form the first initiative of the Mariner Mark 2 program, a series of planetary missions whose common objective is to explore primitive bodies and the outer solar system, toward the ultimate goal of understanding the nature of our origins. Cassini and CRAF are exciting planetary missions. The objectives that they share, the region of the solar system in which comets, asteroids, and the Saturnian system have evolved and now reside, and the spacecraft that will carry both sets of experiments to their targets in the outer solar system are described.

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NASA NTRS1990

Optimal finite-thrust spacecraft trajectories using collocation and nonlinear programming

A new method is described for the determination of optimal spacecraft trajectories in an inverse-square field using finite, fixed thrust. The method employs a recently developed optimization technique which uses a piecewise polynomial representation for the state and controls, and collocation, thus converting the optimal control problem into a nonlinear programming problem, which is solved numerically. This technique has been modified to provide efficient handling of those portions of the trajectory which can be determined analytically, i.e., the coast arcs. Among the problems that have been solved using this method are optimal rendezvous and transfer (including multirevolution cases) and optimal multiburn orbit insertion from hyperbolic approach.

Enright, Paul J.↗

NASA NTRS1993

On-orbit refueling

During the past decade, both NASA and the DOD have conducted numerous space servicing studies. These studies have shown that fluid resupply of on-orbit spacecraft is feasible and would allow for extended spacecraft utilization. In order to prove that the studies have validity, an on-orbit flight demonstration of automatic fluid resupply is required. To embark on this flight demonstration, the system concepts, operational procedures, and conceptual service and target vehicles must be identified. Hernandez Engineering, Inc. (HEI), under the direction of the Space Servicing System Project Office of the NASA/JSC New Initiatives Office, has conducted a systems engineering and integration study. The study objective was to develop preliminary concepts for a flight demonstration of automatic rendezvous, proximity operations, capture, and fluid transfer utilizing servicer and target vehicles. The results show that a servicer vehicle/target kit can be launched to orbit with an ELV and automatically rendezvous and dock with the explorer platform (EP). The servicer vehicle can then separate from the EP/kit, perform proximity maneuvers, redock with the EP/kit, and perform fluid transfer operations. After the on-orbit flight demonstration is completed, the servicer/kit can be separated from the EP and be deorbited into the Earth's atmosphere.

Moore, James S.↗

NASA NTRS1991

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

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.

Tchoryk, Peter, Jr.↗

NASA NTRS2015

BILLIARDS: A Demonstration Mission for Hundred-Meter Class Near-Earth Asteroid Disruption

Collisions from near-Earth asteroids (NEAs) have the potential to cause widespread harm to life on Earth. The hypervelocity nature of these collisions means that a relatively small asteroid (about a quartermile in diameter) could cause a global disaster. Proposed strategies for deflecting or disrupting such a threatening asteroid include detonation of a nuclear explosive device (NED) in close proximity to the asteroid, as well as intercepting the asteroid with a hypervelocity kinetic impactor. NEDs allow for the delivery of large amounts of energy to a NEA for a given mass launched from the Earth, but have not yet been developed or tested for use in deep space. They also present safety and political complications, and therefore may only be used when absolutely necessary. Kinetic impactors require a relatively simple spacecraft compared to NEDs, but also deliver a much lower energy for a given launch mass. To date, no demonstration mission has been conducted for either case, and such a demonstration mission must be conducted prior to the need to utilize them during an actual scenario to ensure that an established, proven system is available for planetary defense when the need arises. One method that has been proposed to deliver a kinetic impactor with impact energy approaching that of an NED is the "billiard-ball" approach. This approach would involve capturing an asteroid approximately ten meters in diameter with a relatively small spacecraft (compared to the launch mass of an equivalent direct kinetic impactor), and redirecting it into the path of an Earth-threatening asteroid. This would cause an impact which would disrupt the Earth-threatening asteroid or deflect it from its Earth-crossing trajectory. The BILLIARDS Project seeks to perform a demonstration of this mission concept in order to establish a protocol that can be used in the event of an impending Earth/asteroid collision. In order to accomplish this objective, the mission must (1) rendezvous with a small (less than 10m), NEA (hereinafter "Alpha"), (2) maneuver Alpha to a collision with a approx. 100 m NEA (hereinafter "Beta"), and (3) produce a detectable deflection or disruption of Beta. In addition to these primary objectives, the BILLIARDS project will contribute to the scientific understanding of the physical properties and collision dynamics of asteroids, and provide opportunities for international collaboration.

Asteroid↗

NASA NTRS1995

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

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.

Source record↗

NASA NTRS1977

Ion Drive multi-asteroid rendezvous trajectories

The general characteristics of multiasteroid rendezvous missions are examined, taking into account a use of an improved solar electric propulsion system, called Ion Drive. The improvements are related to the development of light-weight solar cells, extensions to high thruster power levels with high exhaust velocity, and the attachment of solar concentrators to increase the power available beyond 1.5 AU. Flight times on the order of 500-800 days are generally required. It is assumed that a stay time of about 60 days at each asteroid is adequate. Attention is given to launch and spacecraft parameters, the search technique, and multiasteroid tours involving Ceres and Vesta. It is found that at least four asteroids in the main belt, generally selected from those with known characteristics, may be reached on a single mission with a 100-kg separable probe landed on each one.

Bender, D. F.↗

NASA NTRS1991

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

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.

Tchoryk, Peter, Jr.↗

NASA NTRS2019

Early Navigation Performance of the OSIRIS-REx Approach to Bennu

The New Frontiers-class OSIRIS-REx (Origins, Spectral Interpretation, Resource Identification, Security-Regolith Explorer) mission is the first American endeavor to return a sample from an asteroid. In preparation for retrieving the sample, OSIRIS-REx is conducting a campaign of challenging proximity-operations maneuvers and scientific observations, bringing the spacecraft closer and closer to the surface of near-Earth asteroid (101955) Bennu. Ultimately, the spacecraft will enter a 900-meter-radius orbit about Bennu and conduct a series of reconnaissance flybys of candidate sample sites before being guided into contact with the surface for the Touch and Go sample collection event. Between August and December 2018, the OSIRIS-REx team acquired the first optical observations of Bennu and used them for navigation. We conducted a series of maneuvers with the main engine, Trajectory Correction Maneuver, and Attitude Control System thruster sets to slow the OSIRIS-REx approach to Bennu and achieve rendezvous on December 3, 2018. This paper describes the trajectory design, navigation conops, and key navigation results from the Approach phase of the OSIRIS-REx mission.

Antreasian, Peter G.↗

NASA NTRS2000

Semi-Major Axis Knowledge and GPS Orbit Determination

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.

Carpenter, J. Russell↗

NASA NTRS2000

Semi-Major Axis Knowledge and GPS Orbit Determination

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.

Carpenter, J. Russell↗