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Bearing provides continuous, smooth movement of the cat's-eye mirror, eliminating wear and deterioration of bearing surface and resulting oscillation effects in servo system. Design features self-aligning configuration; single-point, pivotal pad mounting, having air passage through it; and design of pads that allows for precise control of discharge path of air from pads.
In this paper, a new and efficient algorithm is developed for attitude determination from Global Positioning System signals. The new algorithm is derived from a generalized nonlinear predictive filter for nonlinear systems. This uses a one time-step ahead approach to propagate a simple kinematics model for attitude determination. The advantages of the new algorithm over previously developed methods include: it provides optimal attitudes even for coplanar baseline configurations; it guarantees convergence even for poor initial conditions; it is a non-iterative algorithm; and it is computationally efficient. These advantages clearly make the new algorithm well suited to on-board applications. The performance of the new algorithm is tested on a dynamic hardware simulator. Results indicate that the new algorithm accurately estimates the attitude of a moving vehicle, and provides robust attitude estimates even when other methods, such as a linearized least-squares approach, fail due to poor initial starting conditions.
Designing a planetary suit is very complex and often requires difficult trade-offs between performance, cost, mass, and system complexity. In order to verifying that new suit designs meet requirements, full prototypes must eventually be built and tested with human subjects. Using computer models early in the design phase of new hardware development can be advantageous, allowing virtual prototyping to take place. Having easily modifiable models of the suit hard sections may reduce the time it takes to make changes to the hardware designs and then to understand their impact on suit and human performance. A virtual design environment gives designers the ability to think outside the box and exhaust design possibilities before building and testing physical prototypes with human subjects. Reductions in prototyping and testing may eventually reduce development costs. This study is an attempt to develop computer models of the hard components of the suit with known physical characteristics, supplemented with human subject performance data. Objectives: The primary objective was to develop an articulating solid model of the Mark III hip bearings to be used for evaluating suit design performance of the hip joint. Methods: Solid models of a planetary prototype (Mark III) suit s hip bearings and brief section were reverse-engineered from the prototype. The performance of the models was then compared by evaluating the mobility performance differences between the nominal hardware configuration and hardware modifications. This was accomplished by gathering data from specific suited tasks. Subjects performed maximum flexion and abduction tasks while in a nominal suit bearing configuration and in three off-nominal configurations. Performance data for the hip were recorded using state-of-the-art motion capture technology. Results: The results demonstrate that solid models of planetary suit hard segments for use as a performance design tool is feasible. From a general trend perspective, the suited performance trends were comparable between the model and the suited subjects. With the three off-nominal bearing configurations compared to the nominal bearing configurations, human subjects showed decreases in hip flexion of 64%, 6%, and 13% and in hip abduction of 59%, 2%, and 20%. Likewise the solid model showed decreases in hip flexion of 58%, 1%, and 25% and in hip abduction of 56%, 0%, and 30%, under the same condition changes from the nominal configuration. Differences seen between the model predictions and the human subject performance data could be attributed to the model lacking dynamic elements and performing kinematic analysis only, the level of fit of the subjects with the suit, the levels of the subject s suit experience.
A series hybrid system, utilizing a free piston Stirling engine with a linear alternator, and a parallel hybrid system, incorporating a kinematic Stirling engine, are analyzed for various specified reference missions/vehicles ranging from a small two passenger commuter vehicle to a van. Parametric studies for each configuration, detail tradeoff studies to determine engine, battery and system definition, short term energy storage evaluation, and detail life cycle cost studies were performed. Results indicate that the selection of a parallel Stirling engine/electric, hybrid propulsion system can significantly reduce petroleum consumption by 70 percent over present conventional vehicles.
There are few scientific publications comparing human locomotion between motorized and nonmotorized treadmills. Lakomy (1987) and Gamble et al (1988) reported that forward lean is greater on a nonmotorized treadmill to aid in the generation of horizontal force necessary for belt propulsion, but there are no data concerning lower limb kinematics. During long-term spaceflight, astronauts use locomotive exercise to mitigate the physiological effects caused by long-term exposure to microgravity. A critical decision for mission planners concerns the requirements for a treadmill to be used during potential trips to the Moon and Mars. Treadmill operation in an un-powered configuration could reduce mission resource demands, but also may impact the efficacy of treadmill exercise countermeasures. To ascertain the most appropriate type of treadmill to be used, it is important to understand biomechanical differences between motorized (M) and nonmotorized (NM) locomotion. The purpose of this evaluation was to test for differences in lower limb kinematics that occur during M and NM treadmill locomotion at two speeds. It was hypothesized that hip and knee joint angle trajectories would differ between the conditions.
The etalon is shown to remain stable to less than 5-m/sec equivalent wind over one orbit and less than 100 m/sec over several months in orbit. The thermal stability considerations for the highly stable etalon are discussed, including the passive thermal control used in flight. The various physical processes which give rise to drifts in the peak transmission wavelength are quantified, and the kinematic mount used to isolate the etalon from large incremental forces caused by changes in the thermal environment is presented. The performance of the etalon in its flight configuration is evaluated.
A study team at MSFC has conceptually designed two lunar-based telescopes, a 16 m aperture Large Lunar Telescope (LLT) and a 4 m Cluster Telescope Experiment (CTE). The CTE serves as a technology tested for the LLT. The CTE utilizes a hexapod telescope mount to obtain a small gimbal angle that significantly extends the viewing space and observation time from that of a stationary transit telescope. This paper addresses the applicability of six linear actuators to provide limited gimbal rotation of the CTE telescope assembly. Kinematic constraint equations for the hexapod mount have been applied to a six degree-of-freedom rigid body model. Static equilibrium equations have been solved to address configuration design issues, and a feedback tracking control system has been designed and simulated.
Wind tunnel and piloted simulation studies show that the use of the CCV concept of relaxed static pitch stability on fighter aircraft at high angles of attack can result in pitch departures caused by inertia and/or kinematic coupling and in deep stall trim. These problems can be avoided by sufficient nose-down pitch control at high attack angles, and effectiveness of several pitch control configurations such as aft-mounted stabilators, canard mounted flaps, and all-movable canards was investigated. A variable incidence canard with a large travel range was the most effective means of obtaining the required pitch control, although very large deflections may be required on highly unstable configurations. Control laws were developed for cases where the high angle of pitch control requirement is not met, but they can limit aircraft roll capability and thus compromise maneuverability.
Researchers at the Center for Design Research at Stanford University, in collaboration with NASA Ames at Moffet Field, California, are developing hand-powered mechanical prehensors to replace gloves for EVA spacesuits. The design and functional properties of the first version Direct Link Prehensor (DLP) is discussed. It has a total of six degrees-of-freedom and is the most elaborate of three prehensors being developed for the project. The DLP has a robust design and utilizes only linkages and revolute joints for the drive system. With its anthropomorphic configuration of two fingers and a thumb, it is easy to control and is capable of all of the basic prehension patterns such as cylindrical or lateral pinch grasps. Kinematic analysis reveals that, assuming point contacts, a grasped object can be manipulated with three degrees-of-freedom. Yet, in practice more degrees-of-freedom are possible.
Free-piston Stirling power conversion has been considered a candidate for radioisotope power systems for space for more than a decade. Prior to the free-piston Stirling architecture, systems were designed with kinematic Stirling engines with rotary alternators to convert heat to electricity. These systems were proposed with lightly loaded linkages to achieve the necessary life. When the free-piston configuration was initially proposed, it was thought to be attractive due to the relatively high conversion efficiency, acceptable mass, and the potential for long life and high reliability. These features have consistently been recognized by teams that have studied technology options for radioisotope power systems. Since free-piston Stirling power conversion was first considered for space power applications, there have been major advances in three general areas of development: demonstration of life and reliability, the success achieved by Stirling cryocoolers in flight, and the overall developmental maturity of the technology for both flight and terrestrial applications. Based on these advances, free-piston Stirling convertors are currently being developed for a number of terrestrial applications. They commonly operate with the power, efficiency, life, and reliability as intended, and much of the development now centers on system integration. This paper will summarize the accomplishments of free-piston Stirling power conversion technology over the past decade, review the status, and discuss the challenges that remain.
Free-piston Stirling power conversion has been considered a candidate for radioisotope power systems for space for more than a decade. Prior to the free-piston Stirling architecture, systems were designed with kinematic Stirling engines with rotary alternators to convert heat to electricity. These systems were proposed with lightly loaded linkages to achieve the necessary life. When the free-piston configuration was initially proposed, it was thought to be attractive due to the relatively high conversion efficiency, acceptable mass, and the potential for long life and high reliability. These features have consistently been recognized by teams that have studied technology options for radioisotope power systems. Since free-piston Stirling power conversion was first considered for space power applications, there have been major advances in three general areas of development: demonstration of life and reliability, the success achieved by Stirling cryocoolers in flight, and the overall developmental maturity of the technology for both flight and terrestrial applications. Based on these advances, free-piston Stirling convertors are currently being developed for a number of terrestrial applications. They commonly operate with the power, efficiency, life, and reliability as intended, and much of the development now centers on system integration. This paper will summarize the accomplishments of free-piston Stirling power conversion technology over the past decade, review the status, and discuss the challenges that remain.
Vehicle motions resulting from a supersonic mixed-compression inlet unstart were examined to determine if the unstart constituted a hazard severe enough to warrant rejection of mixed-compression inlets on high-speed civil transport (HSCT) concepts. A simple kinematic analysis of an inlet unstart during cruise was performed for a Mach 2, 4, 250-passenger HSCT concept using data from a wind-tunnel test of a representative configuration with unstarted inlets simulated. A survey of previously published research on inlet unstart effects, including simulation and flight test data for the YF-12, XB-70, and Concorde aircraft, was conducted to validate the calculated results. It was concluded that, when countered by suitable automatic propulsion and flight control systems, the vehicle dynamics induced by an inlet unstart are not severe enough to preclude the use of mixed-compression inlets on an HSCT from a passenger safety standpoint. The ability to provide suitable automatic controls appears to be within the current state of the art. However, the passenger startle and discomfort caused by the noise, vibration, and cabin motions associated with an inlet unstart remain a concern.
The MAGIK Robotic analysis team provides kinematic feasibility assessments for the ISS and has produced an animation to evaluate Extravehicular Robotics (EVR) activities for the SpaceX (SpX)-22 flight. The animation also displays the current ISS configuration at the time of the flight. Flight SpX-22 includes a SpaceX Cargo Dragon vehicle docked to Node 2 Zenith (N2Z) International Docking Adapter (IDA) and consists of the delivery of the ISS Power Augmentation (IPA) Carrier. Robotic operations shown include the Space Station Remote Manipulator System (SSRMS) Trunk extraction of the IPA Carrier and maneuvering to the install position on the Payload ORU (Orbital Replacement Unit) Accommodation (POA). There is no disposal payload for this flight.
The MAGIK Robotic analysis team provides kinematic feasibility assessments for the ISS and has produced an animation to evaluate Extravehicular Robotics (EVR) activities for the SpaceX (SpX)-26 flight. The animation also displays the current ISS configuration at the time of the flight. Flight SpX-26 includes a SpaceX Cargo Dragon vehicle docked to Node 2 Zenith (N2Z) International Docking Adapter (IDA) and consists of the delivery of the ISS Roll Out Solar Array (IROSA) mounted to an ISS Power Augmentation (IPA) Carrier. Robotic operations shown include the Space Station Remote Manipulator System (SSRMS) Trunk extraction of the IPA Carrier and maneuvering to the install position on the Payload ORU (Orbital Replacement Unit) Accommodation (POA). There is no disposal payload for this flight.
The MAGIK Robotic analysis team provides kinematic feasibility assessments for the ISS and has produced an animation to evaluate Extravehicular Robotics (EVR) activities for the SpaceX (SpX)-29 flight. The animation also displays the current ISS configuration at the time of the flight. Flight SpX-29 includes a SpaceX Cargo Dragon vehicle docked to Node 2 Forward (N2F) International Docking Adapter (IDA) and consists of the delivery of AWE (Atmospheric Waves Experiment) and Integrated LCRD LEO User Modem Amplifier optical communications Terminal (ILLUMA-T). Robotic operations shown include the Special Purpose Dexterous Manipulator (SPDM) Trunk extraction of ILLUMA-T and maneuvering to the JEM RMS Handoff position, the SPDM Trunk extraction of AWE, maneuvering to Express Logistics Carrier (ELC) 1 Site 3 to remove Space Test Program - Houston 5 (STP-H5), and the installation of AWE to ELC-1 Site 3. There is no disposal payload for this flight.
This paper describes the design and implementation of a real-time control system with multiple modes of operation for a mobile dexterous manipulator. The manipulator under study is a kinematically redundant seven degree-of-freedom arm from Robotics Research Corporation, mounted on a one degree-of-freedom motorized platform.
The Tropical Rainfall Measuring Mission (TRMM) solar arrays are placed into orbital configuration by a complex deployment system. Its two wings each comprise twin seven square solar panels located by a twelve foot articulated boom. The four spring-driven hinge lines per wing are rate-limited by viscous dampers. The wings are stowed against the spacecraft kinematically, and released by five pyrotechnically-actuated mechanisms. Since deployment failure would be catastrophic, a total of 17 deployment tests were completed to qualify the system for the worst cast launch environment. This successful testing culminated in the flawless deployment of the solar arrays on orbit, 15 minutes after launch in November 1997. The custom gravity negation system used to perform deployment testing is modular to allow its setup in several locations, including the launch site in Japan. Both platform and height can be varied, to meet the requirements of the test configuration and the test facility. Its air pad floatation system meets tight packaging requirements, allowing installation while stowed against the spacecraft without breaking any flight interfaces, and avoiding interference during motion. This system was designed concurrently with the deployment system, to facilitate its installation, to aid in the integration of the flight system to the spacecraft, while demonstrating deployment capabilities. Critical parameters for successful testing were alignment of deployment axes and tables to gravity, alignment of table seams to minimize discontinuities, and minimizing pressure drops in the air supply system. Orbital performance was similar to that predicted by ground testing.
Time-dependent fluid behaviors in microgravity environments are examined with particular attention given to the dynamics of bubble configurations. The initial bubble profiles, computed from the steady-state formulation, are used as the initial input for the time dependent computation. The parameters considered are liquid density and its kinematic viscosity, gas density and its kinematic viscosity, surface tension coefficient, angular velocity, and gravity environment. This computer algorithm can be used to simulate the fluid behavior in a microgravity environment, in particular the excitation of slosh waves due to different frequencies of gravity jitters. Results show that lower-frequency gravity jitters excite slosh waves with a higher ratio of maximum amplitude to wave length than that of the slosh waves generated by the higher frequency gravity jitters.