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At least 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Effect of DC Offset on the T-Wave Residuum Parameter

The T-wave residuum (TWR) is a relatively new 12-lead ECG parameter that may reflect cardiac repolarization heterogeneity. TWR shows clinical promise and may become an important diagnostic tool if accurate, consistent, and convenient methods for its calculation can be developed. However, there are discrepancies between the methods that various investigators have used to calculate TWR, as well as some questions about basic methodology and assumptions that require resolution. The presence of a DC offset or very low frequency AC component to the ECG is often observed. Many researchers have attempted to compensate for these by high pass filters and by median beat techniques. These techniques may help minimize the contribution of a low frequency AC component to the TWR, but they will not eliminate a DC offset inherent within the instrumentation. The present study examined the presence of DC offsets in the ECG record, and their effect on TWR. Specifically, in healthy individuals, a DC offset was added to all 8 channels collectively or to each channel selectively. Even with offsets that were relatively small compared to T-wave amplitude, the addition of either collectively or individually applied offsets was observed to produce very significant changes in the TWR, affecting its value by as much as an order of magnitude. These DC offsets may arise from at least two possible sources: a transient artifact from EMG or electrode movement resulting in a transient baseline offset in one or more channels. Since highpass filters have a settling time of several seconds, these artifacts will contribute to a transitory baseline offset lasting 1020 cycles. The machine hardware may also introduce an offset. Regardless of the cause or source of a DC offset, this study demonstrates that offsets have a very significant impact on TWR, and that future studies must not ignore their presence, but rather more appropriately compensate for them.

Scott, N.↗
NASA NTRS2009

International Space Station Environmental Control and Life Support System Acceptance Testing for Node 1 Atmosphere Control and Supply Subsystem

The International Space Station (ISS) Node 1 Environmental Control and Life Support (ECLS) System is comprised of five subsystems: Atmosphere Control and Supply (ACS), Atmosphere Revitalization (AR), Fire Detection and Suppression (FDS), Temperature and Humidity Control (THC), and Water Recovery and Management (WRM). This paper provides a summary of the Node 1 ECLS ACS subsystem design and a detailed discussion of the ISS ECLS Acceptance Testing methodology utilized for that subsystem.

Williams, David E.↗
NASA NTRS2009

Stirling Convertor Control for a Concept Rover at NASA Glenn Research Center

The U.S. Department of Energy (DOE), Lockheed Martin Space Systems Company (LMSSC), Sunpower Inc., and NASA Glenn Research Center (GRC) have been developing an Advanced Stirling Radioisotope Generator (ASRG) for potential use as an electric power system for space science missions. This generator would make use of the free-piston Stirling cycle to achieve higher conversion efficiency than currently used alternatives. NASA GRC initiated an experiment with an ASRG simulator to demonstrate the functionality of a Stirling convertor on a mobile application, such as a rover. The ASRG simulator made use of two Advanced Stirling Convertors to convert thermal energy from a heat source to electricity. The ASRG simulator was designed to incorporate a minimum amount of support equipment, allowing integration onto a rover powered directly by the convertors. Support equipment to provide control was designed including a linear AC regulator controller, constant power controller, and Li-ion battery charger controller. The ASRG simulator is controlled by a linear AC regulator controller. The rover is powered by both a Stirling convertor and Li-ion batteries. A constant power controller enables the Stirling convertor to maintain a constant power output when additional power is supplied by the Li-ion batteries. A Li-ion battery charger controller limits the charging current and cut off current of the batteries. This paper discusses the design, fabrication, and implementation of these three controllers.

Blaze-Dugala, Gina M.↗
NASA NTRS2009

Measuring Multiple Resistances Using Single-Point Excitation

In a proposed method of determining the resistances of individual DC electrical devices connected in a series or parallel string, no attempt would be made to perform direct measurements on individual devices. Instead, (1) the devices would be instrumented by connecting reactive circuit components in parallel and/or in series with the devices, as appropriate; (2) a pulse or AC voltage excitation would be applied at a single point on the string; and (3) the transient or AC steady-state current response of the string would be measured at that point only. Each reactive component(s) associated with each device would be distinct in order to associate a unique time-dependent response with that device.

Hall, Dan↗
NASA NTRS2009

Circuit for Driving Piezoelectric Transducers

The figure schematically depicts an oscillator circuit for driving a piezoelectric transducer to excite vibrations in a mechanical structure. The circuit was designed and built to satisfy application-specific requirements to drive a selected one of 16 such transducers at a regulated amplitude and frequency chosen to optimize the amount of work performed by the transducer and to compensate for both (1) temporal variations of the resonance frequency and damping time of each transducer and (2) initially unknown differences among the resonance frequencies and damping times of different transducers. In other words, the circuit is designed to adjust itself to optimize the performance of whichever transducer is selected at any given time. The basic design concept may be adaptable to other applications that involve the use of piezoelectric transducers in ultrasonic cleaners and other apparatuses in which high-frequency mechanical drives are utilized. This circuit includes three resistor-capacitor networks that, together with the selected piezoelectric transducer, constitute a band-pass filter having a peak response at a frequency of about 2 kHz, which is approximately the resonance frequency of the piezoelectric transducers. Gain for generating oscillations is provided by a power hybrid operational amplifier (U1). A junction field-effect transistor (Q1) in combination with a resistor (R4) is used as a voltage-variable resistor to control the magnitude of the oscillation. The voltage-variable resistor is part of a feedback control loop: Part of the output of the oscillator is rectified and filtered for use as a slow negative feedback to the gate of Q1 to keep the output amplitude constant. The response of this control loop is much slower than 2 kHz and, therefore, does not introduce significant distortion of the oscillator output, which is a fairly clean sine wave. The positive AC feedback needed to sustain oscillations is derived from sampling the current through the piezoelectric transducer. This positive AC feedback, in combination with the slow feedback to the voltage-variable resistors, causes the overall loop gain to be just large enough to keep the oscillator running. The positive feedback loop includes two 16-channel multiplexers, which are not shown in the figure. One multiplexer is used to select the desired piezoelectric transducer. The other multiplexer, which is provided for use in the event that there are significant differences among the damping times of the 16 piezoelectric transducers, facilitates changing the value of one of the resistors in the positive-feedback loop to accommodate the damping time of the selected transducer.

Randall, David P.↗
NASA NTRS2010

Solar Array Disturbances to Spacecraft Pointing During the Lunar Reconnaissance Orbiter (LRO) Mission

The Lunar Reconnaissance Orbiter (LRO), the first spacecraft to support NASA s return to the Moon, launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle. It was initially inserted into a direct trans-lunar trajectory to the Moon. After a five day transit to the Moon, LRO was inserted into the Lunar orbit and successfully lowered to a low altitude elliptical polar orbit for spacecraft commissioning. Successful commissioning was completed in October 2009 when LRO was placed in its near circular mission orbit with an approximate altitude of 50km. LRO will spend at least one year orbiting the Moon, collecting lunar environment science and mapping data, utilizing a suite of seven instruments to enable future human exploration. The objective is to provide key science data necessary to facilitate human return to the Moon as well as identification of opportunities for future science missions. LRO's instrument suite will provide the high resolution imaging data with sub-meter accuracy, highly accurate lunar cartographic maps, mineralogy mapping, amongst other science data of interest. LRO employs a 3-axis stabilized attitude control system (ACS) whose primary control mode, the "Observing Mode", provides Lunar nadir, off-nadir, and inertial fine pointing for the science data collection and instrument calibration. This controller combines the capability of fine pointing with on-demand large angle full-sky attitude reorientation. It provides simplicity of spacecraft operation as well as additional flexibility for science data collection. A conventional suite of ACS components is employed in the Observing Mode to meet the pointing and control objectives. Actuation is provided by a set of four reaction wheels developed in-house at NASA Goddard Space Flight Center (GSFC). Attitude feedback is provided by a six state Kalman filter which utilizes two SELEX Galileo Star Trackers for attitude updates, and a single Honeywell Miniature Inertial Measurement Unit (MIMU) to provide body rates for attitude propagation. Rate is computed by differentiating accumulated angle provided by the MIMU. The Observing Mode controller is required to maintain fine pointing while a large fully-articulated solar array (SA) maintains its panel normal to the solar incidence. This paper describes the disturbances to the attitude control resulting from the SA articulation. Observing Mode performance in the presence of this disturbance was assessed while the spacecraft was in an initial elliptical low altitude orbit during the commissioning phase, which started about two weeks after launch and lasted for 90 days. LRO demonstrated excellent pointing performance during Observing Mode nadir and inertial attitude target operations during this phase. Transient LRO attitude errors observed during commissioning resulted primarily from three sources, Diviner instrument calibrations, RW zero crossings, and SA articulation. Even during times of considerable disturbance from SA articulation, the attitude errors were maintained below the statistical attitude error requirement level of 15 arc-sec (3 sigma).

Calhoun, Philip↗
NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Thruster Control Mode Design and Flight Experience

National Aeronautics and Space Administration s (NASA) Goddard Space Flight Center (GSFC) in Greenbelt, MD, designed, built, tested, and launched the Lunar Reconnaissance Orbiter (LRO) from Cape Canaveral Air Force Station on June 18, 2009. The LRO spacecraft is the first operational spacecraft designed to support NASA s return to the Moon, as part of the Vision for Space Exploration. LRO was launched aboard an Atlas V 401 launch vehicle into a direct insertion trajectory to the Moon. Twenty-four hours after separation the propulsion system was used to perform a mid-course correction maneuver. Four days after the mid-course correction a series of propulsion maneuvers were executed to insert LRO into its commissioning orbit. The commission period lasted eighty days and this followed by a second set of thruster maneuvers that inserted LRO into its mission orbit. To date, the spacecraft has been gathering invaluable data in support of human s future return to the moon. The LRO Attitude Control Systems (ACS) contains two thruster based control modes: Delta-H and Delta-V. The design of the two controllers are similar in that they are both used for 3-axis control of the spacecraft with the Delta-H controller used for momentum management and the Delta-V controller used for orbit adjust and maintenance maneuvers. In addition to the nominal purpose of the thruster modes, the Delta-H controller also has the added capability of performing a large angle slew maneuver. A suite of ACS components are used by the thruster based control modes, for both initialization and control. For initialization purposes, a star tracker or the Kalman Filter solution is used for providing attitude knowledge and upon entrance into the thruster based control modes attitude knowledge is provided via rate propagation using a inertial reference unit (IRU). Rate information for the controller is also supplied by the IRU. Three-axis control of the spacecraft in the thruster modes is provided by eight 5-lbf class attitude control thrusters configured in two sets of four thrusters for redundancy purposes. Four additional 20-lbf class thrusters configured in two sets of two thrusters are used for Lunar Orbit Insertion maneuvers. The propulsion system is one the few systems on-board the LRO spacecraft that has built in redundancy. The Delta-H controller consists of a Proportional-Derivative (PD) controller with a structural filter on the thrusters and a Proportional controller on the reaction wheels. The PD control that employs the thrusters is used for attitude and rate control. The Proportional controller on the reaction wheels is used for commanding the wheels to a new momentum state. The ground commands used for the Delta-H controller are the system momentum vector, reaction wheel momentum, maximum expected command time, and which set of attitude control thrusters to use. The ability to command both the system momentum vector and reaction wheel momentum in the Delta-H controller provides both a capability and an additional source of operator error. Large angle slews via the Delta-H controller is achievable via this commands because these commands are used for the exit mode criteria. Setting these commands to non-consistent values prevents the mode from exiting nominally.

Hsu, Oscar C.↗
NASA NTRS2010

A Modified Lunar Reconnaissance Orbiter (LRO) High Gain Antenna (HGA) Controller Based on Flight Performance

The National Aeronautics and Space Administration's (NASA) Lunar Reconnaissance Orbiter (LRO) was launched on June 18, 2009 and is currently in a 50 km mean altitude polar orbit around the Moon. LRO was designed and built by the NASA Goddard Space Flight Center in Greenbelt, MD. The spacecraft is three-axis stabilized via the attitude control system (ACS), which is composed of various control modes using different sets of sensors and actuators. In addition to pointing the spacecraft, the ACS is responsible for pointing LRO s two appendages, the Solar Array (SA) and the High Gain Antenna (HGA). This study reviews LRO s HGA control system. Starting with an overview of the HGA system, the paper delves into the single input single output (SISO) linear analysis followed by the controller design. Based on flight results, an alternate control scheme is devised to address inherent features in the flight control system. The modified control scheme couples the HGA loop with the spacecraft pointing control loop, and through analysis is shown to be stable and improve transient performance. Although proposed, the LRO project decided against implementing this modification.

Shah, Neerav↗
NASA NTRS2010

Test Rack Development for Extended Operation of Advanced Stirling Convertors at NASA Glenn Research Center

The U.S. Department of Energy, Lockheed Martin Space Systems Company, Sunpower Inc., and NASA Glenn Research Center (GRC) have been developing an Advanced Stirling Radioisotope Generator (ASRG) for use as a power system on space science missions. This generator will make use of free-piston Stirling convertors to achieve higher conversion efficiency than with currently available alternatives. One part of NASA GRC's support of ASRG development includes extended operation testing of Advanced Stirling Convertors (ASCs) developed by Sunpower Inc. and GRC. The ASC consists of a free-piston Stirling engine integrated with a linear alternator. NASA GRC has been building test facilities to support extended operation of the ASCs for several years. Operation of the convertors in the test facility provides convertor performance data over an extended period of time. One part of the test facility is the test rack, which provides a means for data collection, convertor control, and safe operation. Over the years, the test rack requirements have changed. The initial ASC test rack utilized an alternating-current (AC) bus for convertor control; the ASRG Engineering Unit (EU) test rack can operate with AC bus control or with an ASC Control Unit (ACU). A new test rack is being developed to support extended operation of the ASC-E2s with higher standards of documentation, component selection, and assembly practices. This paper discusses the differences among the ASC, ASRG EU, and ASC-E2 test racks.

Dugala, Gina M.↗
NASA NTRS2010

Free-Piston Stirling Power Conversion Unit for Fission Surface Power, Phase I Final Report

This report summarizes the design of a 12 kW dual opposed free-piston Stirling convertor and controller for potential future use in space missions. The convertor is heated via a pumped NaK loop and cooling is provided by a pumped water circuit. Convertor efficiency is projected at 27 percent (AC electrical out/heat in). The controller converts the AC electrical output to 120 Vdc and is projected at 91 percent efficiency. A mechanically simple arrangement, based on proven technology, was selected in which the piston is resonated almost entirely by the working space pressure swing, while the displacer is resonated by planar mechanical springs in the bounce space.

Wood, J. Gary↗
NASA NTRS2008

Multilayer Piezoelectric Stack Actuator Characterization

Future NASA missions are increasingly seeking to use actuators for precision positioning to accuracies of the order of fractions of a nanometer. For this purpose, multilayer piezoelectric stacks are being considered as actuators for driving these precision mechanisms. In this study, sets of commercial PZT stacks were tested in various AC and DC conditions at both nominal and extreme temperatures and voltages. AC signal testing included impedance, capacitance and dielectric loss factor of each actuator as a function of the small-signal driving sinusoidal frequency, and the ambient temperature. DC signal testing includes leakage current and displacement as a function of the applied DC voltage. The applied DC voltage was increased to over eight times the manufacturers' specifications to investigate the correlation between leakage current and breakdown voltage. Resonance characterization as a function of temperature was done over a temperature range of -180C to +200C which generally exceeded the manufacturers' specifications. In order to study the lifetime performance of these stacks, five actuators from one manufacturer were driven by a 60volt, 2 kHz sine-wave for ten billion cycles. The tests were performed using a Lab-View controlled automated data acquisition system that monitored the waveform of the stack electrical current and voltage. The measurements included the displacement, impedance, capacitance and leakage current and the analysis of the experimental results will be presented.

active mirrors↗
NASA NTRS2001

Thermal Analysis for Orbiter and ISS Plume Impingement on International Space Station

The NASA Reaction Control System (RCS) Plume Model (RPM) is an exhaust plume flow field and impingement heating code that has been updated and applied to components of the International Space Station (ISS). The objective of this study was to use this code to determine if plume environments from either Orbiter PRCS jets or ISS reboost and Attitude Control System (ACS) jets cause thermal issues on ISS component surfaces. This impingement analysis becomes increasingly important as the ISS is being assembled with its first permanent crew scheduled to arrive by the end of fall 2000. By early summer 2001 , the ISS will have a number of major components installed such as the Unity (Node 1), Destiny (Lab Module), Zarya (Functional Cargo Block), and Zvezda (Service Module) along with the P6 solar arrays and radiators and the Z-1 truss. Plume heating to these components has been analyzed with the RPM code as well as additional components for missions beyond Flight 6A such as the Propulsion Module (PM), Mobile Servicing System, Space Station Remote Manipulator System, Node 2, and the Cupola. For the past several years NASA/JSC has been developing the methodology to predict plume heating on ISS components. The RPM code is a modified source flow code with capabilities for scarfed nozzles and intersecting plumes that was developed for the 44 Orbiter RCS jets. This code has been validated by comparison with Shuttle Plume Impingement Flight Experiment (SPIFEX) heat flux and pressure data and with CFD and Method of Characteristics solutions. Previous analyses of plume heating predictions to the ISS using RPM have been reported, but did not consider thermal analysis for the components nor jet-firing histories as the Orbiter approaches the ISS docking ports. The RPM code has since been modified to analyze surface temperatures with a lumped mass approach and also uses jet-firing histories to produce pulsed heating rates. In addition, RPM was modified to include plume heating from ISS jets to ISS components where the jet coordinates are specified, together with the engine cant angle. These latter studies have been focused on the PM with plumes from its reboost and ACS jets impinging on various ISS components and also focused on the Japanese H2 Transfer Vehicle (HTV) with the plumes from its reboost engines impinging on the Cupola window. This paper will present plume heating and surface temperature results on a number of ISS components with and without jet-firing histories, evaluate post-flight data, and describe any potential thermal issues

Rochelle, William C.↗
NASA NTRS2011

International Space Station Environmental Control and Life Support System Acceptance Testing for Node 1 Temperature and Humidity Control Subsystem

The International Space Station (ISS) Node 1 Environmental Control and Life Support (ECLS) System is comprised of five subsystems: Atmosphere Control and Storage (ACS), Atmosphere Revitalization (AR), Fire Detection and Suppression (FDS), Temperature and Humidity Control (THC), and Water Recovery and Management (WRM). This paper will provide a summary of the Node 1 ECLS THC subsystem design and a detailed discussion of the ISS ECLS Acceptance Testing methodology utilized for this subsystem.The International Space Station (ISS) Node 1 Environmental Control and Life Support (ECLS) System is comprised of five subsystems: Atmosphere Control and Storage (ACS), Atmosphere Revitalization (AR), Fire Detection and Suppression (FDS), Temperature and Humidity Control (THC), and Water Recovery and Management (WRM). This paper will provide a summary of the Node 1 ECLS THC subsystem design and a detailed discussion of the ISS ECLS Acceptance Testing methodology utilized for this subsystem.

Williams, David E.↗
NASA NTRS2011

Multiplexed Energy Coupler for Rotating Equipment

A multiplexing antenna assembly can efficiently couple AC signal/energy into, or out of, rotating equipment. The unit only passes AC energy while blocking DC energy. Concentric tubes that are sliced into multiple pieces are assembled together so that, when a piece from an outer tube aligns well with an inner tube piece, efficient energy coupling is achieved through a capacitive scheme. With N outer pieces and M inner pieces, an effective N x M combination can be achieved in a multiplexed manner. The energy coupler is non-contact, which is useful if isolation from rotating and stationary parts is required. Additionally, the innovation can operate in high temperatures. Applications include rotating structure sensing, non-contact energy transmission, etc.

Zhao, Xiaoliang↗
NASA NTRS2011

Investigating On-Orbit Attitude Determination Anomalies for the Solar Dynamics Observatory Mission

The Solar Dynamics Observatory (SDO) was launched on February 11, 2010 from Kennedy Space Center on an Atlas V launch vehicle into a geosynchronous transfer orbit. SDO carries a suite of three scientific instruments, whose observations are intended to promote a more complete understanding of the Sun and its effects on the Earth's environment. After a successful launch, separation, and initial Sun acquisition, the launch and flight operations teams dove into a commissioning campaign that included, among other things, checkout and calibration of the fine attitude sensors and checkout of the Kalman filter (KF) and the spacecraft s inertial pointing and science control modes. In addition, initial calibration of the science instruments was also accomplished. During that process of KF and controller checkout, several interesting observations were noticed and investigated. The SDO fine attitude sensors consist of one Adcole Digital Sun Sensor (DSS), two Galileo Avionica (GA) quaternion-output Star Trackers (STs), and three Kearfott Two-Axis Rate Assemblies (hereafter called inertial reference units, or IRUs). Initial checkout of the fine attitude sensors indicated that all sensors appeared to be functioning properly. Initial calibration maneuvers were planned and executed to update scale factors, drift rate biases, and alignments of the IRUs. After updating the IRU parameters, the KF was initialized and quickly reached convergence. Over the next few hours, it became apparent that there was an oscillation in the sensor residuals and the KF estimation of the IRU bias. A concentrated investigation ensued to determine the cause of the oscillations, their effect on mission requirements, and how to mitigate them. The ensuing analysis determined that the oscillations seen were, in fact, due to an oscillation in the IRU biases. The low frequencies of the oscillations passed through the KF, were well within the controller bandwidth, and therefore the spacecraft was actually following the oscillating biases, resulting in movement of the spacecraft on the order of plus or minus 20 arcsec. Though this level of error met the ACS attitude knowledge requirement of [35, 70, 70] arcsec, 3 sigma, the desire of the ACS and instrument teams was to remove as much of the oscillation as possible. The Kearfott IRUs have an internal temperature controller, designed to maintain the IRU temperature at a constant temperature of approximately 70 C, thus minimizing the change in the bias drift and scale factors of the mechanical gyros. During ground testing of the observatory, it was discovered that the 83-Hz control cycle of the IRU heaters put a tremendous amount of stress on the spacecraft battery. Analysis by the power systems team indicated that the constant charge/discharge on the battery due to the IRU thermal control cycle could potentially limit the life of the battery. After much analysis, the decision was made not to run the internal IRU heaters. Analysis of on orbit data revealed that the oscillations in the IRU bias had a connection to the temperature of the IRU; changes in IRU temperature resulted in changes in the amplitude and period of the IRU biases. Several mitigating solutions were investigated, the result of which was to tune the KF with larger IRU noise assumptions which allows the KF to follow and correct for the time-varying IRU biases.

Vess, Melissa F.↗
NASA NTRS2011

Control of Fan Blade Vibrations Using Piezoelectrics and Bi-Directional Telemetry

A novel wireless device which transfers supply power through induction to rotating operational amplifiers and transmits low voltage AC signals to and from a rotating body by way of radio telemetry has been successfully demonstrated in the NASA Glenn Research Center (GRC) Dynamic Spin Test Facility. In the demonstration described herein, a rotating operational amplifier provides controllable AC power to a piezoelectric patch epoxied to the surface of a rotating Ti plate. The amplitude and phase of the sinusoidal voltage command signal, transmitted wirelessly to the amplifier, was tuned to completely suppress the 3rd bending resonant vibration of the plate. The plate's 3rd bending resonance was excited using rotating magnetic bearing excitation while it spun at slow speed in a vacuum chamber. A second patch on the opposite side of the plate was used as a sensor. This paper discusses the characteristics of this novel device, the details of a spin test, results from a preliminary demonstration, and future plans.

Provenza, Andrew J.↗
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 NTRS2011

Instrument Pointing Capabilities: Past, Present, and Future

This paper surveys the instrument pointing capabilities of past, present and future space telescopes and interferometers. As an important aspect of this survey, we present a taxonomy for "apples-to-apples" comparisons of pointing performances. First, pointing errors are defined relative to either an inertial frame or a celestial target. Pointing error can then be further sub-divided into DC, that is, steady state, and AC components. We refer to the magnitude of the DC error relative to the inertial frame as absolute pointing accuracy, and we refer to the magnitude of the DC error relative to a celestial target as relative pointing accuracy. The magnitude of the AC error is referred to as pointing stability. While an AC/DC partition is not new, we leverage previous work by some of the authors to quantitatively clarify and compare varying definitions of jitter and time window averages. With this taxonomy and for sixteen past, present, and future missions, pointing accuracies and stabilities, both required and achieved, are presented. In addition, we describe the attitude control technologies used to and, for future missions, planned to achieve these pointing performances.

celestial targets↗