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NASA NTRS2011

Transient Plume Model Testing Using LADEE Spacecraft Attitude Control System Operations

We have learned it is conceivable that the Neutral Mass Spectrometer on board the Lunarr Atmosphere Dust Environment Explorer (LADEE) could measure gases from surface-reflected Attitude Control System (ACS) thruster plume. At minimum altitude, the measurement would be maximized, and gravitational influence minimized ("short" time-of-flight (TOF) situation) Could use to verify aspects of thruster plume modeling Model the transient disturbance to NMS measurements due to ACS gases reflected from lunar surface Observe evolution of various model characteristics as measured by NMS Species magnitudes, TOF measurements, angular distribution, species separation effects

Woronowicz, Michael↗

NASA NTRS2015

NASA Propulsion Sub-System Concept Studies and Risk Reduction Activities for Resource Prospector Lander

NASA's exploration roadmap is focused on developing technologies and performing precursor missions to advance the state of the art for eventual human missions to Mars. One of the key components of this roadmap is various robotic missions to Near-Earth Objects, the Moon, and Mars to fill in some of the strategic knowledge gaps. The Resource Prospector (RP) project is one of these robotic precursor activities in the roadmap. RP is a multi-center and multi-institution project to investigate the polar regions of the Moon in search of volatiles. The mission is rated Class D and is approximately 10 days, assuming a five day direct Earth to Moon transfer. Because of the mission cost constraint, a trade study of the propulsion concepts was conducted with a focus on available low-cost hardware for reducing cost in development, while technical risk, system mass, and technology advancement requirements were also taken into consideration. The propulsion system for the lander is composed of a braking stage providing a high thrust to match the lander's velocity with the lunar surface and a lander stage performing the final lunar descent. For the braking stage, liquid oxygen (LOX) and liquid methane (LCH4) propulsion systems, derived from the Morpheus experimental lander, and storable bi-propellant systems, including the 4th stage Peacekeeper (PK) propulsion components and Space Shuttle orbital maneuvering engine (OME), and a solid motor were considered for the study. For the lander stage, the trade study included miniaturized Divert Attitude Control System (DACS) thrusters (Missile Defense Agency (MDA) heritage), their enhanced thruster versions, ISE-100 and ISE-5, and commercial-off-the-shelf (COTS) hardware. The lowest cost configuration of using the solid motor and the PK components while meeting the requirements was selected. The reference concept of the lander is shown in Figure 1. In the current reference configuration, the solid stage is the primary provider of delta-V. It will generate 15,000-lbf of thrust with a single burn of ~ 80's seconds. The lander stage is a bi-propellant, pressure-regulated, pulsing liquid propulsion system to perform all other functions.

Trinh, Huu P.↗

NASA NTRS1982

Capability of the manned maneuvering unit to support satellite servicing operations

Features, capabilities, and projected performances and missions of the Manned Maneuvering Unit (MMU) are described. A zero gravity, autonomous, six degree of freedom mobility system using gaseous nitrogen for propulsion, the MMU is equipped with jets for attitude and rotational control, 24 thrusters total, with each providing 7.5 N thrust. The unit is controled from hand controllers used by the pilot, in addition to an automatic attitude control using outputs from a gyroscope pack in the control electronics assembly. EVA periods of 6 hrs are possible with the 150 kg MMU and its fuel complement of 12 kg of gaseous N. Refueling is allowed through an Orbiter interface at the same location the MMU is stowed. The MMU has a delta-V capability of 20 m/sec, with a thrust duration of 7085 N-sec, implying the EVA suit limit will be reached before propellant can be exhausted. The results of various mission simulations for the MMU are reported.

Bergonz, F. H.↗

NASA NTRS2001

Three axis pulsed plasma thruster with angled cathode and anode strip lines

A spacecraft attitude and altitude control system utilizes sets of three pulsed plasma thrusters connected to a single controller. The single controller controls the operation of each thruster in the set. The control of a set of three thrusters in the set makes it possible to provide a component of thrust along any one of three desired axes. This configuration reduces the total weight of a spacecraft since only one controller and its associated electronics is required for each set of thrusters rather than a controller for each thruster. The thrusters are positioned about the spacecraft such that the effect of the thrusters is balanced.

Cassady, R. Joseph↗

NASA NTRS1975

Solutions for discharge chamber sputtering and anode deposit spalling in small mercury ion thrusters

Small electron bombardment mercury ion thrusters are considered for attitude-control and station-keeping operations in the case of earth synchronous satellites. The required operational lifetimes of the order of 20,000 hours could present problems because of the occurrence of sputter erosion of thruster discharge chamber components. The investigation reported is concerned with a study of these problems and the development of approaches for overcoming them.

Power, J. L.↗

NASA NTRS1979

SERT 2 1979 extended flight thruster system performance

Steady state tests of the thruster 2 system on the SERT 2 spacecraft are presented. A direct thrust measurement was obtained for the ion thruster during operations to increase the spacecraft spin rate to maintain spacecraft attitude stability. The continued restart tests of thruster 1 and a report on the general status of all spacecraft systems including the main solar array are presented.

Kerslake, W. R.↗

NASA NTRS1973

Review of Kaufman thruster development at the Lewis Research Center - 1973

Work on Kaufman thruster development completed during the years 1971 and 1972 is reviewed. Thrusters tested have ranged in size from 2.5-cm to 150-cm diameters, in thrust from 0.4 to 4300 mN, and in power from 0.03 to 203 kW. A 2.5-cm thruster was briefly tested and found to have surprisingly high thruster efficiency. Emphasis is placed on thruster system reliability and lifetime as previous work has increased thruster efficiency to a high level. Work also proceeds on definition of thruster-spacecraft interactions. Major R&D efforts are directed at present into two areas of thruster size: a 5-cm to 8-cm diameter thruster to be used for station keeping and attitude control of geosynchronous spacecraft; and a 30-cm diameter thruster to be used for primary propulsion in a 3- to 7-thruster array for solar electric propulsion of interplanetary spacecraft.

Kerslake, W. R.↗

NASA NTRS2008

Attitude Controller for the Atmospheric Entry of the Mars Science Laboratory

This paper describes the attitude controller for the atmospheric entry of the Mars Science Laboratory (MSL). The controller will command 8 RCS thrusters to control the 3- axis attitude of the entry capsule. The Entry Controller is formulated as three independent channels in the control frame, which is nominally aligned with the stability frame. Each channel has a feedfoward and a feedback path. The feedforward path enables fast response to large bank commands. The feedback path stabilizes the vehicle angle of attack and sideslip around its trim position, and tracks bank commands. The feedback path has a PD/D control structure with deadbands that minimizes fuel usage. The performance of this design is demonstrated via computer simulations.

attitude controller↗

NASA NTRS1986

Cosmic Background Explorer (COBE) transfer orbit attitude control system

The Cosmic Background Explorer (COBE) spacecraft will be launched by the Shuttle from Vandenberg AFB into a 300 km altitude, 99 deg inclination, 6 a.m. or 6 p.m. ascending node orbit. After release from the Remote Manipulator System (RMS) arm, an on-board monopropellant hydrazine propulsion system will raise the orbit altitude to 900 km. The spacecraft continuously spins during transfer orbit operations with the spin axis nominally horizontal and in or near the orbit plane. The blowdown propulsion system consists of twelve 5 lb thrusters (3 'spin', 3 'despin', and 6 'axial') with the latter providing initially 30 lb of force parallel to the spin axis for orbit raising. The spin/despin jets provide a constant roll rate during the transfer orbit phase of the mission and the axials control pitch and yaw. The axial thrusters are pulsed on for attitude control during coast periods and are normally on- and off-modulated for control during orbit raising. Attitude sensors employed in the control loops include an array of two-axis digital sun sensors and three planar earth scanners for position measurements, as well as six gyroscopes for rate information. System redundancy is achieved by means of unique three-axes-in-a-plane geometry. This triaxial concept results in a fail-safe operational system with no performance degradation for many different component failure modes.

Placanica, Samuel J.↗

NASA NTRS2005

Attitude Control System Design for the Solar Dynamics Observatory

The Solar Dynamics Observatory mission, part of the Living With a Star program, will place a geosynchronous satellite in orbit to observe the Sun and relay data to a dedicated ground station at all times. SDO remains Sun- pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system is a single-fault tolerant design. Its fully redundant attitude sensor complement includes 16 coarse Sun sensors, a digital Sun sensor, 3 two-axis inertial reference units, 2 star trackers, and 4 guide telescopes. Attitude actuation is performed using 4 reaction wheels and 8 thrusters, and a single main engine nominally provides velocity-change thrust. The attitude control software has five nominal control modes-3 wheel-based modes and 2 thruster-based modes. A wheel-based Safehold running in the attitude control electronics box improves the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integral gains to zero. The paper details the mode designs and their uses.

Starin, Scott R.↗

NASA NTRS2016

Propellantless Attitude Control of Solar Sail Technology Utilizing Reflective Control Devices

Solar sails offer an opportunity for a CubeSatscale, propellant-free spacecraft technology that enables long-term and long-distance missions not possible with traditional methods. Solar sails operate using the transfer of linear momentum from photons of sunlight reflected from the surface of the sail. To propel the spacecraft, no mechanically moving parts, thrusters, or propellant are needed. However, attitude control, or orientation, is still performed using traditional methods involving reaction wheels and propellant ejection, which severely limit mission lifetime. For example, the current state of the art solutions employed by upcoming missions couple solar sails with a state of the art propellant ejection gas system. Here, the use of the gas thruster has limited the lifetime of the mission. To solve the limited mission lifetime problem, the Propellantless Attitude Control of Solar Sail Technology Utilizing Reflective Control Devices project team is working on propellantless attitude control using thin layers of material, an optical film, electrically switchable from transparent to reflective. The technology is based on a polymer-dispersed liquid crystal (PDLC), which allows this switch upon application of a voltage. This technology removes the need for propellant, which reduces weight and cost while improving performance and lifetime.

Solar sail technology↗

NASA NTRS2011

Estimation of Enceladus Plume Density Using Cassini Flight Data

The Cassini spacecraft was launched on October 15, 1997 by a Titan 4B launch vehicle. After an interplanetary cruise of almost seven years, it arrived at Saturn on June 30, 2004. In 2005, Cassini completed three flybys of Enceladus, a small, icy satellite of Saturn. Observations made during these flybys confirmed the existence of water vapor plumes in the south polar region of Enceladus. Five additional low-altitude flybys of Enceladus were successfully executed in 2008-9 to better characterize these watery plumes. During some of these Enceladus flybys, the spacecraft attitude was controlled by a set of three reaction wheels. When the disturbance torque imparted on the spacecraft was predicted to exceed the control authority of the reaction wheels, thrusters were used to control the spacecraft attitude. Using telemetry data of reaction wheel rates or thruster on-times collected from four low-altitude Enceladus flybys (in 2008-10), one can reconstruct the time histories of the Enceladus plume jet density. The 1 sigma uncertainty of the estimated density is 5.9-6.7% (depending on the density estimation methodology employed). These plume density estimates could be used to confirm measurements made by other onboard science instruments and to support the modeling of Enceladus plume jets.

Wang, Eric K.↗

NASA NTRS1990

Space Shuttle vernier thruster long-life chamber development

The Space Shuttle Reaction Control Subsystem (RCS) vernier thruster is a pressure fed engine that utilizes storable propellants to provide precise attitude control for the Orbiter. The current vernier thruster is life limited due to its chamber material. By developing an iridium-lined rhenium chamber for the vernier, substantial gains could be achieved in the operational life of the chamber. The present RCS vernier, its requirements, operating characteristics, and life limitations are described. The current technology status of iridium-lined rhenium is presented along with a description of the operational life capabilities to be gained from implementing this material into the design of a long life vernier chamber. Discussion of the proposed demonstration program to be performed by the NASA Lyndon B. Johnson Space Center to attain additional insight into the application of this technology to the RCS vernier, includes the technical objectives, approach, and program schedule. The plans for further development and integration with the Orbiter and the Shuttle system are also presented.

Krohn, Douglas D.↗

NASA NTRS1971

Design and development of a small structurally integrated ion thruster system

A 5 cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. With 2-D vectorable grids, thrust is 0.56 mlb at a beam voltage of 1200 V, total mass efficiency is 64%, and electrical efficiency is 46.8%. Nonvectoring thrust of 0.41 is demonstrated at very low specific impulse with dielectric coated grids. Structural integrity is demonstrated with dielectric coated grids for shock (30 G), sinusoidal (9 G), and random (19 G squared/Hz) accelerations. System envelope is 31.6 cm long by 14 cm flange B.C. with a mass of 8.5 kg including 6.2 Kg mercury propellant.

Hyman, J., Jr.↗

NASA NTRS1973

Performance optimized, small structurally integrated ion thruster system

A 5-cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. As optimized with a conventional ion extraction system, the system demonstrates a thrust T = 0.47 mlb at a beam voltage of 1600 V, total mass efficiency of 76%, and electrical efficiency of 56%. Under the subject contract effort, no significant performance change was noted for operation with two dimensional electrostatic thrust-vectoring grids. Structural integrity with the vectoring grids was demonstrated for shock (+ or - 30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.2 cm long by 13.4 cm flange bolt circle, with a mass of 9.0 Kg, including 6.8 Kg mercury propellant.

Hyman, J., Jr.↗

NASA NTRS1973

Electric propulsion status and development plans - NASA programs

The current status of the NASA auxiliary and primary electric propulsion programs is reviewed. Plans for continuation of these programs and future target goals are also outlined. In auxiliary propulsion, progress on the ion engine flight experiment for the Application Technology Satellite (ATS)-F launch in 1974 to demonstrate North-South station keeping capability is reviewed. In addition, work on small pulsed plasma thrusters and hydrazine resistojets for attitude control applications are also described. In primary propulsion, the effort aimed at demonstrating the technology for solar electric propulsion is presented.

Lazar, J.↗

NASA NTRS1975

Study of liquid oxygen/liquid hydrogen auxiliary propulsion systems for the space tug

Design concepts are considered that permit use of a liquid-liquid (as opposed to gas-gas) oxygen/hydrogen thrust chamber for attitude control and auxiliary propulsion thrusters on the space tug. The best of the auxiliary propulsion system concepts are defined and their principal characteristics, including cost as well as operational capabilities, are established. Design requirements for each of the major components of the systems, including thrusters, are developed at the conceptual level. The competitive concepts considered use both dedicated (separate tanks) and integrated (propellant from main propulsion tanks) propellant supply. The integrated concept is selected as best for the space tug after comparative evaluation against both cryogenic and storable propellant dedicated systems. A preliminary design of the selected system is established and recommendations for supporting research and technology to further the concept are presented.

Nichols, J. F.↗

NASA NTRS1981

Engineering aspects of the Stanford relativity gyro experiment

According to certain theoretical predictions, the Newtonian laws of motion must be corrected for the effect of a gravitational field. Schiff (1960) proposed an experiment which would demonstrate the effect predicted by Einstein's Theory of General Relativity on a gyroscope. The experiment has been under development at Stanford University since 1961. The requirements involved make it necessary that the test be performed in a satellite to take advantage of weightlessness in space. In a discussion of engineering developments related to the experiment, attention is given to the development of proportional helium thrusters, the simulation of the attitude control system, aspects of inner loop control, the mechanization of the two-loop attitude control system, the effects of helium slosh on spacecraft pointing, and the data instrumentation system.

Everitt, C. W. F.↗