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Improved joystick allows motion in 6 degrees of freedom, biased toward central position and orientation by 16 segments of cable serving as springs. Improvement in feel and control results from nonlinear compliance of cable-spring assembly. Nonlinear variations accommodate natural reactions of hand and brain. Operator functions as part of feedback control loop. More comfortable, increases ability to exert control and reduces fatigue.
The Space Active Vibration Isolation (SAVI) is a concept for vibration isolation of one body from another with simultaneous precise control in 6 Degrees Of Freedom (DOF). SAVI achieves this using a combination of electromechanical linear actuators and magnetic actuators. Other mechanisms of interest include a structure for simulating the body being pointed, an apparatus to simulate the body that is the vibration source, and mechanisms to off-load the weight of each of these two bodies from the experiment to approximate a zero-g condition. A SAVI was built and tested to demonstrate these capabilities.
The Cargo Transfer Vehicle (CTV) will be required to perform six degree of freedom (6 DOF) maneuvers while carrying a wide range of payloads varying from 100,000 lbm to no payload. The current baseline design configuration for the CTV uses a forward propulsion module (FPM) mounted in front of the payload with the CTV behind the payload so that the center of gravity (CG) of the combined stack is centered between the thruster sets. This allows for efficient rotations and translations of heavy payloads in all directions; however, the FPM is a costly item, so it is desirable to find design solutions that do not require the FPM. This presentation provides an overview of the analysis of the FPM requirements for the CTV. In this study, only the reaction control system (RCS) thruster configurations are considered for 6 DOF maneuvers of various CTV cargo configurations. An important output of this study are the viable alternative thruster configurations that eliminate the need for the FPM. Initial results were derived using analytical techniques and simulation analysis tools. Results from the preliminary analysis were validated using our 6 DOF simulation.
Shuttle to Space station docking has become an important issue in the last few years. Docking sensors have been proposed that will provide high precision measurements required for the fuel efficient rendezvous and docking of space vehicles. These sensors will also be used for satellite servicing and orbital assembly. The performance of the docking sensors must be tested before they are implemented in a space environment. A Six-Degree-of-Freedom (6-DOF) Test Facility has been developed at the Tracking & Communications Section, Johnson Space Center to test the static and dynamic accuracies of docking sensors. A candidate sensor is evaluated by comparing the sensor's static position and velocity measurements to the more accurate 6-DOF system. The 6-DOF Test Facility's hardware is fully integrated. Software has been developed in-house to support system operation. The system has been tested statically and the operational parameters verified. System accuracies remain to be determined. Dynamic testing of the facility is expected to begin shortly. Several companies such as McDonnell Douglas, Autonomous Technologies, and General Dynamics are scheduled to test sensors in the next few months. The 6-DOF facility will be available for use in November 1991. This presentation describes the subcomponents, operation, and capabilities of the 6-DOF Test Facility. Discussions will be held on system accuracies. Additional applications of the 6-DOF system will also be addressed.
Proposed future space exploration, such as lunar and Martian expeditions, will require autonomous docking of space vehicles. One proposed candidate method of autonomous docking utilizes a actively controlled parallel manipulator. Operation of the proposed docking manipulator can be segmented into four successive events: prealignment, capture/latching, attenuation, and structural rigidization. This paper discusses the development and testing of a digitally controlled, six-degree-of-freedom (6-DOF), parallel manipulator for the prealignment segment of a docking spacecraft.
The purpose of this study is to determine the adequacy of the Global Positioning System (GPS) in providing relative navigation for automated rendezvous and proximity operations. The study was performed using the Proximity Operations Simulator (POS), Lockheed's high-fidelity, 6 degree of freedom simulation of the space shuttle and space station. This simulation includes identical models of GPS receivers for each vehicle. The navigation software in each vehicle includes identical Kalman filters. Each filter computes the absolute state of its vehicle, and the relative state vector is obtained by simply subtracting absolute states.
The NASA Johnson Space Center (JSC) is involved in the development of an autonomous docking ground demonstration. The demonstration combines the technologies, expertise and facilities of the JSC Tracking and Communications Division (EE), Structures and Mechanics Division (ES), and the Navigation, Guidance and Control Division (EG) and their supporting contractors. The autonomous docking ground demonstration is an evaluation of the capabilities of the laser sensor system to support the docking phase (12ft to contact) when operated in conjunction with the Guidance, Navigation and Control Software. The docking mechanism being used was developed for the Apollo Soyuz Test Program. This demonstration will be conducted using the Six-Degrees of Freedom (6-DOF) Dynamic Test System (DTS). The DTS environment simulates the Space Station Freedom as the stationary or target vehicle and the Orbiter as the active or chase vehicle. For this demonstration the laser sensor will be mounted on the target vehicle and the retroreflectors on the chase vehicle. This arrangement was used to prevent potential damage to the laser. The sensor system. GN&C and 6-DOF DTS will be operated closed-loop. Initial condition to simulate vehicle misalignments, translational and rotational, will be introduced within the constraints of the systems involved. Detailed description of each of the demonstration components (e.g., Sensor System, GN&C, 6-DOF DTS and supporting computer configuration) including their capabilities and limitations will be discussed. A demonstration architecture drawing and photographs of the test configuration will be presented.
TRW has conducted an extensive Contact Dynamics Test Program (CDTP) of the Three Point Docking Mechanism (TPDM). The CDTP tested the ability of the TPDM latches to capture and automatically dock to target spacecraft. The target selected was the Hubble Space Telescope (HST). Mock ups of the TPDM with its three latches and the docking interface of the HST were constructed at the Marshall Space Flight Center (MSFC) in Huntsville, Alabama for use in the tests. The tests were performed at the Flat Floor and Six Degree of Freedom (6-DOF) facilities at MSFC.
In modern fighter aircraft air-to-air missile launch acceptable ranges (LARs) are provided to the pilot for his launch decision. These LARs are calculated in the avionic software, using very simple algebraic equations and algorithms, that attempt to model complex missile behavior. Frequently these algorithms are revised to reflect inaccuracies discovered, when comparisons are made with missile six degree-of-freedom (6-DOF) simulations. The process of modifying this software has become expensive. In the past, these simple LAR generators were modified based only on subjective judgmental approach, with little regard to the missile range design limits. With reduced budgets a more rational methodology should be taken to assess whether or not simple fighter LAR generators need to be modified. Any LAR generator can produce large amounts of range (minimum and maximum) data over the board engagement limits of modem fighter and air-to-air missiles. This infinite population lends itself to some statistical methodology to access simple LAR accuracy and the need for software modification. The methodology should produce statistical parameters for use in the decision making process of modifying a simple fighter LAR generator. This paper presents a regression method, based on a missile range design limit philosophy. Acceptance criteria from the regression method are given that provide simple statistical parameters to aid in this decision making process.
We developed a force-reflecting teleoperation system that uses a real-time graphic simulator. This system eliminates the effects of communication time delays in remote robot manipulation. The simulator provides the operator with predictive display and feedback of computed contact forces through a six-degree of freedom (6-DOF) master arm on a real-time basis. With this system, peg-in-hole tasks involving round-trip communication time delays of up to a few seconds were performed at three support levels: a real image alone, a predictive display with a real image, and a real-time graphic simulator with computed-contact-force reflection and a predictive display. The experimental results indicate the best teleoperation efficiency was achieved by using the force-reflecting simulator with two images. The shortest work time, lowest sensor maximum, and a 100 percent success rate were obtained. These results demonstrate the effectiveness of simulated-force-reflecting teleoperation efficiency.
The present docking system for the Orbiter uses mechanical capture latches that are actuated by contact forces. The forces are generated when the two approaching masses collide at the docking mechanism. There is always a trade-off between having high enough momentum to effect capture and low enough momentum to avoid structural overload or unacceptable angular displacements. The use of the present docking system includes a contact thrusting maneuver that causes high docking loads to be included into Space Station. A magnetic docking aid has been developed to reduce the load s during docking. The magnetic docking aid is comprised of two extendible booms that are attached adjacent to the docking structure with electromagnets attached on the end of the boom. On the mating vehicle, two steel plates are attached. As the Orbiter approaches Space Station, the booms are extended, and the magnets attach to the actuated (without thrusting), by slowly driving the extendible booms to the stowed position, thus reacting the load into the booms. This results in a docking event that has lower loads induced into Space Station structure. This method also greatly simplifies the Station berthing tasks, since the Shuttle Remote Manipulation System (SRMS) arm need only place the element to be berthed on the magnets (no load required), rather than firing the Reaction Control System (RCS) jets to provide the required force for capture latch actuation. The Magnetic Docking Aid was development testing on a six degree-of-freedom (6 DOF) system at JSC.
JPL and Microdexterity Systems collaborated to develop new surgical capabilities. They developed a Robot Assisted Microsurgery (RAM) tool for surgeons to use for operating on the eye, ear, brain, and blood vessels with unprecedented dexterity. A surgeon can hold the surgical instrument with motions of 6 degrees of freedom with an accuracy of 25 microns in a 70 cu cm workspace. In 1996 a demonstration was performed to remove a microscopic particle from a simulated eyeball. In 1997, tests were performed at UCLA to compare telerobotics with mechanical operations. In 5 out of 7 tests, the RAM tool performed with a significant improvement of preciseness over mechanical operation. New design features include: (1) amplified forced feedback; (2) simultaneous slave robot instrumentation; (3) index control switch on master handle; and (4) tool control switches. Upgrades include: (1) increase in computational power; and (2) installation of hard disk memory storage device for independent operation and independent operation of forceps. In 1997 a final demonstration was performed using 2 telerobotics simultaneously in a microsurgery suture procedure to close a slit in a thin sheet of latex rubber which extended the capabilities of microsurgery procedures. After completing trials and demonstrations for the FDA the potential benefits for thousands of operations will be exposed.
An initial assessment of a proposed High-Speed Civil Transport (HSCT) was conducted in the fall of 1995 at the NASA Langley Research Center. This configuration, known as the Industry Reference-H (Ref.-H), was designed by the Boeing Aircraft Company as part of their work in the High Speed Research program. It included a conventional tail, a cranked-arrow wing, four mixed-flow turbofan engines, and capacity for transporting approximately 300 passengers. The purpose of this assessment was to evaluate and quantify operational aspects of the Reference-H configuration from a pilot's perspective with the additional goal of identifying design strengths as well as any potential configuration deficiencies. This study was aimed at evaluating the Ref.-H configuration at many points of the aircraft's envelope to determine the suitability of the vehicle to accomplish typical mission profiles as well as emergency or envelope-limit conditions. Pilot-provided Cooper-Harper ratings and comments constituted the primary vehicle evaluation metric. The analysis included simulated real-time piloted evaluations, performed in a 6 degree of freedom motion base NASA Langley Visual-Motion Simulator, combined with extensive bath analysis. The assessment was performed using the third major release of the simulation data base (known as Ref.-H cycle 2B).
NASA Dryden Flight Research Center uses its six-degrees-of-freedom (6-DOF) fixed-base simulations for mission control room training to improve flight safety and operations. This concept is applied to numerous flight projects such as the F-18 High Alpha Research Vehicle (HARV), the F-15 Intelligent Flight Control System (IFCS), the X-38 Actuator Control Test (XACT), and X-43A (Hyper-X). The Dryden 6-DOF simulations are typically used through various stages of a project, from design to ground tests. The roles of these simulations have expanded to support control room training, reinforcing flight safety by building control room staff proficiency. Real-time telemetry, radar, and video data are generated from flight vehicle simulation models. These data are used to drive the control room displays. Nominal static values are used to complete information where appropriate. Audio communication is also an integral part of training sessions. This simulation capability is used to train control room personnel and flight crew for nominal missions and emergency situations. Such training sessions are also opportunities to refine flight cards and control room display pages, exercise emergency procedures, and practice control room setup for the day of flight. This paper describes this technology as it is used in the X-43A and F-15 IFCS and XACT projects.
The Space Technology-7 Disturbance Reduction System is being designed to demonstrate the ability to shield a test mass from non-gravitational forces. In order to meet this goal, two advanced technologies will be employed: a highly sensitive Gravitational Reference Sensor and micro-Newton thrusters. ST-7 is limited to two clusters of four thrusters, which are sufficient to provide control for 6 degrees of freedom, but the overall effectiveness of the baseline configuration is limited by the noise on the measurement signals and the thrust outputs, the disturbance force caused by solar radiation pressure, and the performance of the individual thrusters. This paper presents and discusses these issues in greater detail along with possible mitigation methods.
The Mars Microprobe mission will provide the first opportunity for subsurface measurements, including water detection, near the south pole of Mars. In this paper, performance of the Microprobe aeroshell design is evaluated through development of a six-degree-of-freedom (6-DOF) aerodynamic database and flight dynamics simulation. Numerous mission uncertainties are quantified and a Monte-Carlo analysis is performed to statistically assess mission performance. Results from this 6-DOF Monte-Carlo simulation demonstrate that, in a majority of the cases (approximately 2-sigma), the penetrator impact conditions are within current design tolerances. Several trajectories are identified in which the current set of impact requirements are not satisfied. From these cases, critical design parameters are highlighted and additional system requirements are suggested. In particular, a relatively large angle-of-attack range near peak heating is identified.
A generic control system framework for both real-time and batch six-degree-of-freedom (6-DOF) simulations is presented. This framework uses a simplified dynamic inversion technique to allow for stabilization and control of any type of aircraft at the pilot interface level. The simulation, designed primarily for the real-time simulation environment, also can be run in a batch mode through a simple guidance interface. Direct vehicle-state acceleration feedback is required with the simplified dynamic inversion technique. The estimation of surface effectiveness within real-time simulation timing constraints also is required. The generic framework provides easily modifiable control variables, allowing flexibility in the variables that the pilot commands. A direct control allocation scheme is used to command aircraft effectors. Primary uses for this system include conceptual and preliminary design of aircraft, when vehicle models are rapidly changing and knowledge of vehicle 6-DOF performance is required. A simulated airbreathing hypersonic vehicle and simulated high-performance fighter aircraft are used to demonstrate the flexibility and utility of the control system.
To control microgravity environment on the International Space Station (ISS), NASA developed payloads have to meet the payload integration requirements of the Space Station Program, specifically a microgravity allocation plan. The Microgravity Emissions Laboratory (MEL) was developed at NASA Glenn Research Center (GRC) for verification of the payloads compliance with payload integration requirements. MEL is a 6 degree of freedom inertial measurement system capable of characterizing the microgravity emissions, generated by a disturber, down to a micro g. Microgravity Emissions tests provide a payload developer with a tool to assess payload's compliance with the requirements, i.e. forces and moments, generated by the payload at its center of gravity. Forces and moments are presented in time domain for both stationary and transient signals, and in frequency domain for the stationary signals. To date, MEL conducted over thirty tests of ISS hardware. The test results are being successfully used by the payload developers for design verification and improvement.