Search NASA⌕ Search

SEARCH · Search NASA

Results for “Reaction Control System Thrusters”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Dawn Spacecraft Reaction Control System Flight Experience

The NASA Dawn spacecraft mission is studying conditions and processes of the solar system's earliest epoch by investigating two protoplanets remaining intact since their formations, Ceres and Vesta. Launch was in 2007. Ion propulsion is used to fly to and enter orbit around Vesta, depart Vesta and fly to Ceres, and enter orbit around Ceres. A conventional blowdown hydrazine reaction control system (RCS) is used to provide external torques for attitude control. Reaction wheel assemblies were intended to provide attitude control in most cases. However, the spacecraft experienced one, then two apparent failures of reaction wheels. Also, similar thrusters experienced degradation in a long life application on another spacecraft. Those factors led to RCS being operated in ways completely different than anticipated prior to launch. Numerous mitigations and developments needed to be implemented. The Vesta mission was fully successful. Even with the compromises necessary due to those anomalies, the Ceres mission is also projected to be feasible.

Mizukami, Masashi↗
NASA NTRSDate not supplied

NASA Gateway Refueling Architecture and Concept of Operation

The Lunar Gateway is a deep space orbiting outpost being developed by NASA in partnership with ESA and other domestic and international partners. It is a critical component of NASA’s Artemis program supporting long-term human exploration of the moon and is designed to be refueled, requiring the on-orbit transfer of propellants. The first two modules of Gateway to be launched will be the Power and Propulsion Element (PPE) and the Habitation and Logistics Outpost (HALO). The PPE contains the bipropellant chemical and electrical propulsion systems that will provide attitude control and orbit raising capability for Gateway. The reaction control system (RCS) utilizes monomethylhydrazine (MMH) and mixed oxides of nitrogen-3 (MON-3) as the fuel and oxidizer, respectively. Helium is used as a pressurant. The Advanced Electric Propulsion Systems (AEPS) and Busek hall thrusters that comprise the solar electric propulsion (SEP) system use xenon as the propellent. The ESPRIT Refueling Module, provided by ESA, will supply the propellant refueling function to Gateway’s propulsion systems. On-orbit refueling is a complex, technically challenging operation that is key to enabling sustainable crewed Lunar and Martian exploration. Numerous systems may be involved in the refueling operation. Also numerous systems are required to enable the successful ability for Gateway to refuel such as, SEP, RCS, Structure & Mechanisms (S&M), Guidance, Navigation and Control (GNC), Thermal, Software, Vehicle System Manager (VSM), Flight Ops, Extra Vehicular Robotics (EVR), Communication and Tracking (C&T), Avionics, and more. This paper will detail the design architecture, concept of operations, and challenges associated with the Gateway refueling system. The methodologies used in Gateway will be compared to those implemented in OSAM-1 and the best practices documented by AIAA and CONFERS (Consortium For Execution Of Rendezvous And Servicing Operations).

Brandie L. Rhodes↗
NASA NTRS2007

Fault Detection and Correction for the Solar Dynamics Observatory Attitude Control System

The Solar Dynamics Observatory is an Explorer-class mission that will launch in early 2009. The spacecraft will operate in a geosynchronous orbit, sending data 24 hours a day to a devoted ground station in White Sands, New Mexico. It will carry a suite of instruments designed to observe the Sun in multiple wavelengths at unprecedented resolution. The Atmospheric Imaging Assembly includes four telescopes with focal plane CCDs that can image the full solar disk in four different visible wavelengths. The Extreme-ultraviolet Variability Experiment will collect time-correlated data on the activity of the Sun's corona. The Helioseismic and Magnetic Imager will enable study of pressure waves moving through the body of the Sun. The attitude control system on Solar Dynamics Observatory is responsible for four main phases of activity. The physical safety of the spacecraft after separation must be guaranteed. Fine attitude determination and control must be sufficient for instrument calibration maneuvers. The mission science mode requires 2-arcsecond control according to error signals provided by guide telescopes on the Atmospheric Imaging Assembly, one of the three instruments to be carried. Lastly, accurate execution of linear and angular momentum changes to the spacecraft must be provided for momentum management and orbit maintenance. In th~sp aper, single-fault tolerant fault detection and correction of the Solar Dynamics Observatory attitude control system is described. The attitude control hardware suite for the mission is catalogued, with special attention to redundancy at the hardware level. Four reaction wheels are used where any three are satisfactory. Four pairs of redundant thrusters are employed for orbit change maneuvers and momentum management. Three two-axis gyroscopes provide full redundancy for rate sensing. A digital Sun sensor and two autonomous star trackers provide two-out-of-three redundancy for fine attitude determination. The use of software to maximize chances of recovery from any hardware or software fault is detailed. A generic fault detection and correction software structure is used, allowing additions, deletions, and adjustments to fault detection and correction rules. This software structure is fed by in-line fault tests that are also able to take appropriate actions to avoid corruption of the data stream.

Starin, Scott R.↗
NASA NTRS2014

A Comparison of Thruster Implementation Strategies for a Deep Space Nanosatellite

Attitude control for very small spacecraft, commonly referred to as nanosatellites or CubeSats, has traditionally been carried out using reaction wheels and magnetic torquers as the primary actuators. However, as these spacecraft begin to be considered for a broader range of scientific applications, including some beyond low Earth orbit, it has become necessary to also consider thruster systems for actuation. In recent years a number of thruster designs that conform to the mass, volume, and power constraints of nanosatellites have become commercially available, including cold gas systems, pulsed plasma thrusters (PPTs), and micro-electrospray propulsion (MEP) systems. The challenge now facing the nanosatellite community is to determine which thruster solutions are appropriate for a particular application, and what the best method of control might be. This paper will compare the implementation of a cold gas system with that of an MEP or PPT system for an upcoming nanosatellite mission using a previously reported saturation-restricted control law. Results are presented for this controller both with and without a fuel-optimal thruster allocation scheme, and an assessment on incorporating these technologies in an upcoming NASA mission is offered.

Nehrenz, Matt↗
NASA NTRS2019

Method for Deriving Telescope Specifications for Earth-Detecting Coronagraphs and Its Use in Prioritizing Technology Development Investments

Different potential exoplanet mission concepts require different enabling and enhancing technologies. Decisions regarding prioritizing the development of these technologies can be made using a Science-Driven Systems-Engineering framework. The white paper makes the following specific recommendations. (1) Any potential telescope/coronagraph combination should develop a wavefront stability error budget using the method described in this white paper. That error budget will inform discussions about what technology development will be needed to enable the mission. (2) Enhancing and enabling technologies requiring investment include: low noise reaction wheels or control moment gyros; micro-thruster technology; active vibration isolation; materials and design practice for making lightweight stiff structures and optical components; 200-hertz wavefront sensing and control; 500-hertz sub-picometer-precision edge sensor and position actuators; non-destructive process to quantify CTE (Coefficient of Thermal Expansion) homogeneity of 4-meter class mirrors to an uncertainty of plus or minus 1 parts per billion per degree Kelvin over a 100 by 100 sampling; laser gauge positional metrology system; and predictive active thermal control.

Space Mirror Technology↗
NASA NTRS2009

Trade Study of Multiple Thruster Options for the Mars Airplane Concept

A trade study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction (PARR) project (2004-2005) to examine the option of using multiple, smaller thrusters in place of a single large thruster on the Mars airplane concept with the goal to reduce overall cost, schedule, and technical risk. The 5-lbf (22N) thruster is a common reaction control thruster on many satellites. Thousands of these types of thrusters have been built and flown on numerous programs, including MILSTAR and Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion system and compared the results to those of the baseline single thruster system.

Kuhl, Christopher A.↗
NASA NTRS2013

Next-Generation Maneuvering System with Control-Moment Gyroscopes for Extravehicular Activities Near Low-Gravity Objects

Looking ahead to the human exploration of Mars, NASA is planning for exploration of near-Earth asteroids and the Martian moons. Performing tasks near the surface of such low-gravity objects will likely require the use of an updated version of the Manned Maneuvering Unit (MMU) since the surface gravity is not high enough to allow astronauts to walk, or have sufficient resistance to counter reaction forces and torques during movements. The extravehicular activity (EVA) Jetpack device currently under development is based on the Simplified Aid for EVA Rescue (SAFER) unit and has maneuvering capabilities to assist EVA astronauts with their tasks. This maneuvering unit has gas thrusters for attitude control and translation. When EVA astronauts are performing tasks that require ne motor control such as sample collection and equipment placement, the current control system will re thrusters to compensate for the resulting changes in center-of-mass location and moments of inertia, adversely affecting task performance. The proposed design of a next-generation maneuvering and stability system incorporates control concepts optimized to support astronaut tasks and adds control-moment gyroscopes (CMGs) to the current Jetpack system. This design aims to reduce fuel consumption, as well as improve task performance for astronauts by providing a sti er work platform. The high-level control architecture for an EVA maneuvering system using both thrusters and CMGs considers an initial assessment of tasks to be performed by an astronaut and an evaluation of the corresponding human-system dynamics. For a scenario in which the astronaut orbits an asteroid, simulation results from the current EVA maneuvering system are compared to those from a simulation of the same system augmented with CMGs, demonstrating that the forces and torques on an astronaut can be significantly reduced with the new control system actuation while conserving onboard fuel.

Carpenter, Michele↗
NASA NTRS1978

Attitude control of large solar power satellites

Satellite power systems are a promising future source of electrical energy. However, the very large size solar power satellites (relative to contemporary spacecraft) requires investigation of the resulting attitude control problems and of appropriate control techniques. The principal effects of the large size are a great increase in sensitivity to gravity-gradient torques and a great reduction in structural bending frequencies with the attendant likelihood of undesirable control system interaction. A wide variety of control techniques are investigated to define approaches that minimize implementation penalties. These techniques include space-constructed momentum wheels, gravity-gradient stabilization, quasi-inertial free-drift modes, and various reaction control thruster types, some of which reduce the implementation penalties to a few percent of the spacecraft mass. The control system/structural dynamic interaction problem is found to have a tractable solution. Some of the results can be applied to other large space structure spacecraft.

Oglevie, R. E.↗
NASA NTRS1987

Force and moment tests to determine the interaction effects of the reaction control system jet plumes on the space shuttle Orbiter aerodynamics at Mach Number 6 (Test OA352)

The purpose of this test was to expand the existing Space Shuttle aerodynamics and Reaction Control System (RCS) data base to support the Glide Return to Launch Site (GRTLS) abort trajectory and the new Digital Autopilot. An existing model of the orbiter was used to investigate the aerodynamic effects of several combinations of RCS thrusters and thruster momentum ratios at Mach number 6. Two separate model installations were used to achieve an angle-of-attack range of -11 to 46 deg. The test was conducted at a unit Reynolds number of 0.8 x 10 to the 6th per foot.

Cayse, Robert W.↗
NASA NTRS1987

Control systems for autonomous operation of the Magellan spacecraft

The Magellan low cost interplanetary mission spacecraft, scheduled for flight in 1989, employs a computer-based Attitude and Articulation Control System (AACS) that interfaces with spacecraft sensors and actuators. The sensors include two attitude reference units, a star-scanner unit, two digital sun sensors, and a tachometer. The actuators encompass three reaction wheels, a total of 12 attitude control thrusters, eight attitude-control engines, a solar array drive assembly, and an input-output drive assembly. A detailed account is given of the AACS's onboard computer architecture.

Huang, Whittak H.↗
NASA NTRS1987

Independent Orbiter Assessment (IOA): Analysis of the reaction control system, volume 2

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results for the Reaction Control System (RCS). The RCS is situated in three independent modules, one forward in the orbiter nose and one in each OMS/RCS pod. Each RCS module consists of the following subsystems: Helium Pressurization Subsystem; Propellant Storage and Distribution Subsystem; Thruster Subsystem; and Electrical Power Distribution and Control Subsystem. Volume 2 continues the presentation of IOA analysis worksheets.

Burkemper, V. J.↗
NASA NTRS1987

Independent Orbiter Assessment (IOA): Analysis of the reaction control system, volume 3

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results for the Reaction Control System (RCS). The RCS is situated in three independent modules, one forward in the orbiter nose and one in each OMS/RCS pod. Each RCS module consists of the following subsystems: Helium Pressurization Subsystem; Propellant Storage and Distribution Subsystem; Thruster Subsystem; and Electrical Power Distribution and Control Subsystem. Volume 3 continues the presentation of IOA analysis worksheets and the potential critical items list.

Burkemper, V. J.↗
NASA NTRSDate not supplied

Feasibility and Utility of a Cryogenic Integrated RCS

Traditional in-space propulsion systems use storable propellants, such as MMH/NTO, for applications beyond Low Earth Orbit (LEO). Modern advancements in Cryogenic Fluid Management (CFM) technologies opens the door for use of more efficient cryogenic propellants, such as LOX/LCH4 or LOX/LH2, in these long-duration missions beyond LEO. As these Main Propulsion Systems (MPS) transition to cryogenic propellants, a cryogenic solution for Reaction Control Systems (RCS) becomes attractive for several reasons. A mixed-fluid vehicle solution with cryogenic (cool) MPS and storable (warm) RCS complicates thermal management. Additionally, a vehicle with common cryogenic propellant across the MPS and RCS could take advantage of shared hardware such as tanks, thermal management equipment, and pressurization system, reducing mass and development time. The fluid system design proposed in this presentation features shared hardware, leading to the designation of Cryogenic Integrated Reaction Control System (iRCS). The iRCS design stores cryogenic propellant in the low-pressure (10’s psia) MPS tanks, creates steady high-pressure (100’s psia) flow through an electric pump (e-pump), feeds pulse-firing RCS thrusters, and returns any excess flow to the tank through a pressure control device on a recirculation line. This iRCS design is inspired by the automotive fuel-rail, where similar elevated-pressure and pulse-firing requirements are levied. This design has several inherent advantages. Because this design is pump-fed rather than pressure-fed, heavy, high-pressure propellant storage is not required. The RCS propellant can therefore be stored by increasing MPS tank volume. Furthermore, by increasing the pressure of the circulated fluid through an e-pump, the propellant quality moves further into the sub-cooled regime. Sub-cooled liquid is desirable for accurate pressure and flow control at the thruster inlet. Lastly, the iRCS allows the distribution system hardware to maintain cryogenic temperatures by slowly recirculating liquid in a low-energy “Idle Mode”. Altogether, the iRCS concept enables a more efficient cryogenic vehicle design.

Integrated↗
NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Guidance, Navigation and Control (GN&C) Overview

The National Aeronautics and Space Administration s (NASA) Lunar Reconnaissance Orbiter (LRO) launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle and into a direct insertion trajectory to the oon. LRO, which was designed, built, and operated by the NASA Goddard Space Flight Center in Greenbelt, MD, is gathering crucial data on the lunar environment that will help astronauts prepare for long-duration lunar expeditions. The mission has a nominal life of 1 year as its seven instruments find safe landing sites, locate potential resources, characterize the radiation environment, and test new technology. To date, LRO has been operating well within the bounds of its requirements and has been collecting excellent science data images taken from the LRO Camera Narrow Angle Camera of the Apollo landing sites appeared on cable news networks. A significant amount of information on LRO s science instruments is provided at the LRO mission webpage. LRO s Guidance, Navigation and Control (GN&C) subsystem is made up of an onboard attitude control system (ACS) and a hardware suite of sensors and actuators. The LRO onboard ACS is a collection of algorithms based on high level and derived requirements, and reflect the science and operational events throughout the mission lifetime. The primary control mode is the Observing mode, which maintains the lunar pointing orientation and any offset pointing from this baseline. It is within this mode that all science instrument calibrations, slews and science data is collected. Because of a high accuracy requirement for knowledge and pointing, the Observing mode makes use of star tracker (ST) measurement data to determine an instantaneous attitude pointing. But even the star trackers alone do not meet the tight requirements, so a six-state Kalman Filter is employed to improve the noisy measurement data. The Observing mode obtains its rate information from an inertial reference unit (IRU) and in the event of an IRU failure, the rate data is be derived from the star tracker, but with degraded pointing performance. The Delta-V control mode responsibility is to maintain attitude pointing during the cruise trajectory, insertion burns and lunar orbit maintenance by adjustments made to the spacecraft s velocity magnitude and vector direction. The ACS also provides for a thruster based system momentum management algorithm (known as Delta-H) to maintain the system and wheel momentum to within acceptable levels. In the event an anomaly causes the LRO spacecraft to lose the ability to maintain its current attitude pointing, a Sun Safe mode is included in the ACS for the purpose of providing a known power and thermally safe coarse inertial sun attitude for an indefinite period of time, within the manageable limits of the reaction wheels. The Sun Safe mode is also the initial spacecraft control mode off of the launch vehicle and provides for a means to null tip-off rates immediately after separation. The nominal configuration is to use the IRU for rate information in the controller. In the event of a gyro failure a gyroless control mode was developed that computes rate information from the CSS data.

Garrick, Joseph↗
NASA NTRS1994

Basic features of the STS/Spacelab vibration environment

The Space Shuttle acceleration environment is characterized. The acceleration environment is composed of a residual or quasi-steady component and higher frequency components induced by vehicle structural modes and the operation of onboard machinery. Quasi-steady accelerations are generally due to atmospheric drag, gravity gradient effects, and rotational forces. These accelerations tend to vary with the orbital frequency (approx. 10(exp -4) Hz) and have magnitudes less than or equal to 10(exp -6) g(sub 0) (where 1 g(sub 0) is terrestrial gravity). Higher frequency g-jitter is characterized by oscillatory disturbances in the 1-100 Hz range and transient components. Oscillatory accelerations are related to the response of large flexible structures like antennae, the Spacelab module, and the Orbiter itself, and to the operation of rotating machinery. The Orbiter structural modes in the 1-10 Hz range, are excited by oscillatory and transient disturbances and tend to dominate the energy spectrum of the acceleration environment. A comparison of the acceleration measurements from different Space Shuttle missions reveals the characteristic signature of the structural modes of the Orbiter overlaid with mission specific hardware induced disturbances and their harmonics. Transient accelerations are usually attributed to crew activity and Orbiter thruster operations. During crew sleep periods, the acceleration levels are typically on the order of 10(exp -6) g(sub 0) (1 micro-g). Crew work and exercise tend to raise the accelerations to the 10(exp -3) g(sub 0) (1 milli-g) level. Vernier reaction control system firings tend to cause accelerations of 10(exp -4) g(sub 0), while primary reaction control system and Orbiter maneuvering system firings cause accelerations as large as 10(exp -2) g(sub 0). Vibration isolation techniques (both active and passive systems) used during crew exercise have been shown to significantly reduce the acceleration magnitudes.

Baugher, Charles R.↗
NASA NTRS1994

Modal Test Technology as Non-Destructive Evaluation of Space Shuttle Structures

Modal test and analysis Is being used for nondestructive evaluation of Space Shuttle structures. The purpose of modal testing is to measure the dynamic characteristics of a structure to extract its resonance frequencies, damping, and mode shapes. These characteristics are later compared to subsequently acquired characteristics. Changes in the modal characteristics indicate damage in the structure. Use of modal test technology as a damage detection tool was developed at JSC during the Shuttle acoustic certification program and subsequent test programs. The Shuttle Modal Inspection System was created in order to inspect areas that are impossible or impractical to inspect with conventional methods. Areas on which this technique has been applied include control surfaces, which are covered with thermal protection tiles, and the Forward Reaction Control Module, which is a frame structure that supports various tanks, thrusters, and fluid lines, which requires major disassembly to inspect. This paper traces the development of the technology, gives a status of its implementation on the Shuttle, explains challenges involved in implementing this type of inspection program, and suggests future improvements in data analysis and interpretation. Dual-use applications of the technology include inspections of bridges, oil-platforms, and aircraft.

Micheal S Grygler↗
NASA NTRS2002

A Novel Approach for Modeling Chemical Reaction in Generalized Fluid System Simulation Program

The Generalized Fluid System Simulation Program (GFSSP) is a computer code developed at NASA Marshall Space Flight Center for analyzing steady state and transient flow rates, pressures, temperatures, and concentrations in a complex flow network. The code, which performs system level simulation, can handle compressible and incompressible flows as well as phase change and mixture thermodynamics. Thermodynamic and thermophysical property programs, GASP, WASP and GASPAK provide the necessary data for fluids such as helium, methane, neon, nitrogen, carbon monoxide, oxygen, argon, carbon dioxide, fluorine, hydrogen, water, a hydrogen, isobutane, butane, deuterium, ethane, ethylene, hydrogen sulfide, krypton, propane, xenon, several refrigerants, nitrogen trifluoride and ammonia. The program which was developed out of need for an easy to use system level simulation tool for complex flow networks, has been used for the following purposes to name a few: Space Shuttle Main Engine (SSME) High Pressure Oxidizer Turbopump Secondary Flow Circuits, Axial Thrust Balance of the Fastrac Engine Turbopump, Pressurized Propellant Feed System for the Propulsion Test Article at Stennis Space Center, X-34 Main Propulsion System, X-33 Reaction Control System and Thermal Protection System, and International Space Station Environmental Control and Life Support System design. There has been an increasing demand for implementing a combustion simulation capability into GFSSP in order to increase its system level simulation capability of a liquid rocket propulsion system starting from the propellant tanks up to the thruster nozzle for spacecraft as well as launch vehicles. The present work was undertaken for addressing this need. The chemical equilibrium equations derived from the second law of thermodynamics and the energy conservation equation derived from the first law of thermodynamics are solved simultaneously by a Newton-Raphson method. The numerical scheme was implemented as a User Subroutine in GFSSP.

Sozen, Mehmet↗
NASA NTRS1992

Spacecraft attitude and velocity control system

A spacecraft attitude and/or velocity control system includes a controller which responds to at least attitude errors to produce command signals representing a force vector F and a torque vector T, each having three orthogonal components, which represent the forces and torques which are to be generated by the thrusters. The thrusters may include magnetic torquer or reaction wheels. Six difference equations are generated, three having the form ##EQU1## where a.sub.j is the maximum torque which the j.sup.th thruster can produce, b.sub.j is the maximum force which the j.sup.th thruster can produce, and .alpha..sub.j is a variable representing the throttling factor of the j.sup.th thruster, which may range from zero to unity. The six equations are summed to produce a single scalar equation relating variables .alpha..sub.j to a performance index Z: ##EQU2## Those values of .alpha. which maximize the value of Z are determined by a method for solving linear equations, such as a linear programming method. The Simplex method may be used. The values of .alpha..sub.j are applied to control the corresponding thrusters.

Paluszek, Michael A.↗