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The applications, design, control and testing of an actuator that provides the precise motion control of a gimbal platform without torquing against the basebody to which it is attached are described. The reactionless actuator described was given the name reactuator.
The Pointed Telescope Subassembly (PTS) consists of a telescope, its Coude relay optics, a two axis gimbal mechanism, and a cover/caging device. These components, their mechanisms, the requirements, and some of the trade-offs leading to the final design are described. The PTS supplies light to the interferometer of the High Resolution Doppler Imager to be used to study upper atmospheric wind velocity.
The design and in-orbit performance data are presented for the Stratospheric Aerosol and Gas Experiment II (SAGE II) instrument which was launched by Shuttle on the Earth Radiation Budget Satellite. SAGE II is a sun photometer that measures the extinction of solar radiation caused by the earth's atmosphere in seven spectral channels ranging in center wavelength from 0.385 to 1.02 micrometers. These measurements, which occur twice each orbit during satellite sunrise and sunset, are inverted to yield vertical distributions of stratospheric aerosols, ozone, water vapor, and nitrogen dioxide. The SAGE II instrument consists of a Cassegrain telescope with a two axis gimbal mounting, a grating spectrometer, and a 12 bit data system. The instrument tracks the solar centroid in the azimuth plane and vertically scans the instrument's instantaneous field of view across the sun for tangent altitudes ranging from the earth's horizon to 150 km. SAGE II is a third generation instrument following the highly successful Stratospheric Aerosol Measurement II (SAM II) and SAGE I programs.
A softmounted inertially reacting pointing system differs from traditional gimbal-based pointing system architecture in that: (1) the primary pointing control actuator does not need to apply torques on the basebody, and hence will not interact in a destabilizing way with basebody flexibility, and (2) the connection of the payload with the basebody is via a soft, low frequency structure, which acts as a two-way low pass filter for disturbances. Planar, linear analysis of a preliminary design of such a pointing system using the piezoelectric polymer material PVF2 as an active element in the softmount is presented demonstrating the potential performance in disturbance rich environments such as Space Station.
The development and testing of a computer control system for a two-engine xenon ion propulsion module is described. The computer system controls all aspects of the propulsion module operation including: start-up, steady-state operation, throttling and shutdown of the engines; start-up, operation and shutdown of the central neutralizer subsystem; control of the gimbal system for each engine; and operation of the valves in the propellant storage and distribution system. The most important engine control algorithms are described in detail. These control algorithms provide flexibility in the operation and throttling of ion engines which has never before been possible. This flexibility is made possible in large part through the use of flow controllers which maintain the total flow rate of propellant into the engine at the proper level. Data demonstrating the throttle capabilities of the engine and control system are presented.
The high speed, high resolution optical tracking system for the Spaceborne Geodynamic Ranging System employs a two-axis gimbaled pointing device that can operate from a Space Shuttle platform and can track multiple retroreflector ground targets with arcsec accuracy. Laboratory tests of the stepping characteristics of the pointing system for various step sizes and directions has shown arcsec repeatability with little wasted motion, overshoot, or ringing. The worst rms tracking jitter was 1 and 2 arcsec in the roll and pitch axes, respectively, at the maximum tracking rate of 2 deg/sec.
The Space Station polar platform will carry a variety of earth observation instruments for NASA and the National Oceanic and Atmospheric Administration. In this paper, the asymmetrical platform is modeled as three connected rigid bodies. A generalized angular momentum equation is employed to derive the rotational equations of motion. These equations are linearized and used for preliminary sizing of control devices using a classical control approach. Two control systems are considered to stabilize the platform and satisfy the pointing requirements. The first system is composed of a single variable-speed, double-gimbaled momentum wheel and the second consists of three-reaction wheels. The performance of each system is assessed using a linear optimal control approach.
The HIRIS is designed to acquire images in 196 spectral bands simultaneously in the 0.4-2.5 micron wavelength region. HIRIS is a targeting rather than a continuous acquisition instrument and obtains high spatial and spectral resolution images in a 23 km swath with a 30 m GIFOV in vertical viewing. Gimbal pointing is proposed which will allow image acquisition at -30 + 60 deg down-track and + or - 25 deg cross-track. The raw data rate of the instrument is 393 Mbs. The high spectral resolution will make it possible to directly identify surficial materials such as rocks, soils, and suspended matter in water, and HIRIS opens up the possibility of studying biogeochemical processes in vegetation canopies. HIRIS will be used in conjunction with MODIS as a multistage sampling system.
Equations of motion for a three-bladed hingeless rotor coupled to a gimbal-supported fuselage are developed. A brief derivation of the equations is presented, including the assumptions made, the ordering scheme, and a discussion of both the source of the aerodynamic axial load term and the reason for differences due to hinge sequence. The solution of the equations is discussed, including linearization, the use of multiblade coordinates to eliminate periodic coefficients, and specialization to a root-sprung model. The treatment of the blade root offset by using an ordering scheme to neglect higher-order terms is found to affect results in pitch and roll. A small aerodynamic loading term in the blade axial direction has a significant effect at large collective pitch settings, as does the choice of a flap-lag or a lag-flap transformation sequence for blade motion. The results for a hingeless rotor also differ significantly from those for the root-sprung model when the nonrotating natural frequencies are matched.
Hughes Aircraft Company developed qualified and integrated flight, a flight test Ion Auxiliary Propulsion System (IAPS), on an Air Force technology satellite. The IAPS Flight Package consists of two identical Thruster Subsystems and a Diagnostic Subsystem. Each thruster subsystem (TSS) is comprised of an 8-cm ion Thruster-Gimbal-Beam Shield Unit (TGBSU); Power Electronics Unit; Digital Controller and Interface Unit (DCIU); and Propellant Tank, Valve and Feed Unit (PTVFU) plus the requisite cables. The Diagnostic Subsystem (DSS) includes four types of sensors for measuring the effect of the ion thrusters on the spacecraft and the surrounding plasma. Flight qualifications of IAPS, prior to installation on the spacecraft, consisted of performance, vibration and thermal-vacuum testing at the unit level, and thermal-vacuum testing at the subsystem level. Mutual compatibility between IAPS and the host spacecraft was demonstrated during a series of performance and environmental tests after the IAPS Flight Package was installed on the spacecraft. After a spacecraft acoustic test, performance of the ion thrusters was reverified by removing the TGBSUs for a thorough performance test at Hughes Research Laboratories (HRL). The TGBSUs were then reinstalled on the spacecraft. The IAPS Flight Package is ready for flight testing when Shuttle flights are resumed.
Viewgraphs and charts from a briefing summarize the accomplishments, results, conclusions, and recommendations of a feasibility study using the Pinhole Occulter Facility (POF). Accomplishments for 1986 include: (1) improved IPS Gimbal Model; (2) improved Crew Motion Disturbance Model; (3) use of existing shuttle on-orbit simulation to study the effects of orbiter attitude deadband size on POF performance; (4) increased understanding of maximum performance expected from current actuator/sensor set; (5) use of TREETOPS nonlinear time domain program to obtain system dynamics describing the complex multibody flexible structures; (6) use of HONEY-X design tool to design and evaluate multivariable compensator for stability, robustness, and performance; (7) application of state-of-the-art compensator design methodology Linear Quadratic Gaussian/Loop Transfer Recovery (LQG/LTR); and (8) examination of tolerance required on knowledge of the POF boom flexible mode frequencies to insure stability, using structure uncertainty analysis.
A conceptual design study of an aeroassisted orbital transfer vehicle is discussed. Nicknamed TAXI, it will ferry personnel and cargo: (1) between low Earth orbit and a spacecraft circling around the Sun in permanent orbit intersecting gravitational fields of Earth and Mars, and (2) between the cycling spacecraft and a Mars orbiting station, co-orbiting with Phobos. Crew safety and mission flexibility (in terms of ability to provide a wide range of delta-V) were given high priority. Three versions were considered, using the same overall configuration based on a low L/D aerobrake with the geometry of a raked off elliptical cone with ellipsoidal nose and a toroidal skirt. The propulsion system consists of three gimballed LOX/LH2 engines firing away from the aerobrake. The versions differ mainly in the size of the aeroshields and propellant tanks. TAXI A version resulted from an initial effort to design a single transfer vehicle able to meet all delta-V requirements during the 15-year period (2025 to 2040) of Mars mission operations. TAXI B is designed to function with the cycling spacecraft moving in a simplified, nominal trajectory. On Mars missions, TAXI B would be able to meet the requirements of all the missions with a relative approach velocity near Mars of less than 9.3 km/sec. Finally, TAXI C is a revision of TAXI A, a transfer vehicle designed for missions with a relative velocity near Mars larger than 9.3 km/sec. All versions carry a crew of 9 (11 with modifications) and a cargo of 10000 lbm. Trip duration varies from 1 day for transfer from LEO to the cycling ship to nearly 5 days for transfer from the ship to the Phobos orbit.
The study and revision of the gimbal design of the Three-Meter Balloon Borne Telescope (TMBBT) is discussed. Efforts were made to eliminate the alignment and limited rotation problems inherent in the flex-pivot design. A new design using ball bearings to replace the flex-pivots was designed and its performance analyzed. An error analysis for the entire gondola pointing system was also prepared.
Two and three dimensional internal flowfield solutions for critical points in the Space Shuttle solid rocket booster burn time were developed using the Lockheed Huntsville GIM/PAID Navier-Stokes solvers. These perfect gas, viscous solutions for the high performance motor characterize the flow in the aft segment and nozzle of the booster. Two dimensional axisymmetric solutions were developed at t = 20 and t = 85 sec motor burn times. The t = 85 sec solution indicates that the aft segment forward inhibitor stub produces vortices with are shed and convected downwards. A three dimensional 3.5 deg gimbaled nozzle flowfield solution was developed for the aft segment and nozzle at t = 9 sec motor burn time. This perfect gas, viscous analysis, provided a steady state solution for the core region and the flow through the nozzle, but indicated that unsteady flow exists in the region under the nozzle nose and near the flexible boot and nozzle/case joint. The flow in the nozzle/case joint region is characterized by low magnitude pressure waves which travel in the circumferential direction. From the two and three dimensional flowfield calculations presented it can be concluded that there is no evidence from these results that steady state gas dynamics is the primary mechanism resulting in the nozzle pocketing erosion experienced on SRM nozzles 8A or 17B. The steady state flowfield results indicate pocketing erosion is not directly initiated by a steady state gas dynamics phenomenon.
The approach and results of a Life Cycle Cost (LCC) analysis of the Space Station Solar Dynamic Power Subsystem (SDPS) including gimbal pointing and power output performance are documented. The Multi-Discipline Design Tool (MDDT) computer program developed during the 1986 study has been modified to include the design, performance, and cost algorithms for the SDPS as described. As with the Space Station structural and control subsystems, the LCC of the SDPS can be computed within the MDDT program as a function of the engineering design variables. Two simple examples of MDDT's capability to evaluate cost sensitivity and design based on LCC are included. MDDT was designed to accept NASA's IMAT computer program data as input so that IMAT's detailed structural and controls design capability can be assessed with expected system LCC as computed by MDDT. No changes to IMAT were required. Detailed knowledge of IMAT is not required to perform the LCC analyses as the interface with IMAT is noninteractive.
This report describes how guide stars are selected for the Optical Sensor Package (OSP) for the Instrument Pointing System (IPS) when it is operating in the stellar mode on the ASTRO missions. It also describes how the objective loads are written and how the various roll angles are related; i.e., the celestial roll or position angle, the objective load roll angles, and the IPS gimbal angles. There is a brief description of how the IPS operates and its various modes of operation; i.e., IDOP, IDIN, and OSPCAL.
The boresight alignment testing for the sun sensor assembly on the Halogen Occultation Experiment (HALOE) is described. The sun sensor assembly consists of three sensors that provide feedback signals for controlling dual axes gimbals. Two energy balancing silicon detectors are operated as wideband sensors in the azimuth and elevation axes. The third sensor is a silicon photodiode array operated as a narrow-band sensor in the elevation axis. These sensors are mounted on a common Invar structure which is mounted to the HALOE telescope. A blackbody was used as the stimulating source to perform the initial boresight alignment and this was checked with a heliostat solar look and a direct solar look. These tests are explained with a comparison between each source used.
The method of active modal damping (AMD) is reviewed, and the pinhole/occulter facility (P/OF) is presented as a design example. This system is a large space system composed of a flexible beam, a gimbal-pointing system, and an optical alignment system mounted in the Shuttle cargo bay and excited by typical Shuttle disturbances. The AMD system performance is compared with that of a series-compensated control system.