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Conceptual system controls aircraft with old and new guidance equipment. Report begins with overview of concept, then reviews controller-interactive simulations. Describes fuel-conservative-trajectory algorithm, based on equations of motion for controlling landing time. Finally, presents results of piloted simulations.
An overview is given of methodology and test results pertaining to the characterization of ultra sensitive accelerometers. Two issues are of primary concern. The terminology ultra sensitive accelerometer is used to imply instruments whose noise floors and resolution are at the state of the art. Hence, the typical approach of verifying an instrument's performance by measuring it with a yet higher quality instrument (or standard) is not practical. Secondly, it is difficult to find or create an environment with sufficiently low background acceleration. The typical laboratory acceleration levels will be at several orders of magnitude above the noise floor of the most sensitive accelerometers. Furthermore, this background must be treated as unknown since the best instrument available is the one to be tested. A test methodology was developed in which two or more like instruments are subjected to the same but unknown background acceleration. Appropriately selected spectral analysis techniques were used to separate the sensors' output spectra into coherent components and incoherent components. The coherent part corresponds to the background acceleration being measured by the sensors being tested. The incoherent part is attributed to sensor noise and data acquisition and processing noise. The method works well for estimating noise floors that are 40 to 50 dB below the motion applied to the test accelerometers. The accelerometers being tested are intended for use as feedback sensors in a system to actively stabilize an inertial guidance component test platform.
Development of the Open Path Tunable Infrared Monitor of the Atmosphere (OPTIMA) instrument involved designing a pair of motion systems that could maintain a precise alignment and spatial distance between two mirrors installed on the NASA DC-8 research laboratory aircraft. This is the first airborne optical instrument that allows direct measurement of the gases in the freestream airflow on the exterior of the aircraft. One mirror is mounted within a specially constructed open port cavity in the cabin of the aircraft and the second is mounted 6 meters away on top of the inboard port side (number 2) engine pylon. Three co-aligned laser beams are reflected between the two mirrors 64 times in a Herriott pattern. The resulting sample path length of 384 meters is used to perform a spectral absorption analysis of the airflow between the mirrors. To compensate for normal wing movement and engine oscillations both mirrors were designed as continuously driven mechanisms to maintain alignment within allowable limits. The motion systems of the two mirror assemblies provide five degrees of freedom and are designed to maintain a pointing accuracy within seven arc-sec with a response frequency in excess of 10 Hz. The pylon motion system incorporates controlled pitch and yaw movement. The fuselage motion system compensates for pitch variation as well as linear translation for focal length and vertical aiming of the laser beam via a controlled beam guidance mechanism.
Development of the Open Path Tunable Infrared Monitor of the Atmosphere (OPTIMA) instrument involved designing a pair of motion systems that could maintain a precise alignment and spatial distance between two mirrors installed on the NASA DC-8 research laboratory aircraft. This is the first airborne optical instrument that allows direct measurement of the gases in the freestream airflow on the exterior of the aircraft. One mirror is mounted within a specially constructed open port cavity in the cabin of the aircraft and the second is mounted 6 meters away on top of the inboard port side (number 2) engine pylon. Three co-aligned laser beams are reflected between the two mirrors 64 times in a Herriott pattern. The resulting sample path length of 384 meters is used to perform a spectral absorption analysis of the airflow between the mirrors. To compensate for normal wing movement and engine oscillations both mirrors were designed as continuously driven mechanisms to maintain alignment within allowable limits. The motion systems of the two mirror assemblies provide five degrees of freedom and are designed to maintain a pointing accuracy within seven arc-sec with a response frequency in 6xcess of 10 Hz. The pylon motion system incorporates controlled pitch and yaw movement. The fuselage motion system compensates for pitch variation as well as linear translation for focal length and vertical aiming of the laser beam via a controlled beam guidance mechanism.
NASA's Orion spacecraft is designed to autonomously rendezvous and dock with many vehicles including the International Space Station. However, the crew is able to assume manual control of the vehicle s attitude and flight path. In these instances, Orion must meet handling qualities requirements established by NASA. Two handling qualities assessments were conducted at the Johnson Space Center to evaluate preliminary designs of the vehicle using a six degree of freedom, high-fidelity guidance, navigation, and control simulation. The first assessed Orion s handling qualities during the last 20 ft before docking, and included both steady and oscillatory motions of the docking target. The second focused on manual acquisition of the docking axis during the proximity operations phase and subsequent station-keeping. Cooper-Harper handling qualities ratings, workload ratings and comments were provided by 10 evaluation pilots for the docking study and 5 evaluation pilots for the proximity operations study. For the docking task, both cases received 90% Level 1 (satisfactory) handling qualities ratings, exceeding NASA s requirement. All ratings for the ProxOps task were Level 1. These evaluations indicate that Orion is on course to meet NASA's handling quality requirements for ProxOps and docking.
Free-flying models are commonly used f o r path planning and open loop control design (i. e., guidance design) and translational feedback control design (i. e., control design) for deep space precision formation flying. The free flying model, essentially a double integrator, results from discarding small terms in the relative spacecraft equations of motion. While the magnitude of these discarded terms may be small, one must show that their dynamic effects are small as compared to the precision performance requirements. We do so by deriving a theoretical method for bounding the difference between the solution of a nonlinear truth model of the relative translational spacecraft dynamics and a Simplified linear time-invariant model. Presently, the method incorporates feedforward and static output feedback control. The method is applied to a Terrestrial Planet Finder- based example. Using only feedforward control (guidance) the free-flying model and a Hill- Clohessy- Wiltshire Equations-based model are shown to be accurate to 1 c m for up to 4 and 30 hours, respectively. Also shown is that the simplest free-flying model may not be sufficient for low-gain feedback control design-closed-loop tracking errors can be as large as 8 meters.
Adaptation to research paradigms such as rotating rooms and optical alteration of visual feedback during movement results in development of perceptual-motor programs that provide the reflexive assistance that is necessary to skilled control of movement and balance. The discomfort and stomach awareness that occur during the adaptation process has been attributed to conflicting sensory information about the state of motion. Vestibular signals depend on the kinematics of head movements irrespective of the presence or absence of signals from other senses. We propose that sensory conflict when vestibular signals are at least one component of the conflict are innately disturbing and unpleasant. This innate reaction is part of a continuum that operates early in life to prevent development of inefficient perceptual-motor programs. This reaction operates irrespective of and in addition to reward and punishment from parental guidance or goal attainment to yield efficient control of whole body movement in the operating environment of the individual. The same mechanism is involved in adapting the spatial orientation system to strange environments. This conceptual model "explains" why motion sickness is associated with adaptation to novel environments and is in general consistent with motion sickness literature.
Future missions to Mars may require pin-point landing precision, possibly on the order of tens of meters. The ability to reach a target while meeting a dynamic pressure constraint to ensure safe parachute deployment is complicated at Mars by low atmospheric density, high atmospheric uncertainty, and the desire to employ only bank angle control. The vehicle aerodynamic performance requirements and guidance necessary for 0.5 to 1.5 lift drag ratio vehicle to maximize the achievable footprint while meeting the constraints are examined. A parametric study of the various factors related to entry vehicle performance in the Mars environment is undertaken to develop general vehicle aerodynamic design requirements. The combination of low lift drag ratio and low atmospheric density at Mars result in a large phugoid motion involving the dynamic pressure which complicates trajectory control. Vehicle ballistic coefficient is demonstrated to be the predominant characteristic affecting final dynamic pressure. Additionally, a speed brake is shown to be ineffective at reducing the final dynamic pressure. An adaptive precision entry atmospheric guidance scheme is presented. The guidance uses a numeric predictor-corrector algorithm to control downrange, an azimuth controller to govern crossrange, and analytic control law to reduce the final dynamic pressure. Guidance performance is tested against a variety of dispersions, and the results from selected tests are presented. Precision entry using bank angle control only is demonstrated to be feasible at Mars.
This paper proposes a systematic approach for identifying the perturbation parameter in singular perturbation analysis of aircraft optimal guidance, and in particular considers a family of problems related to aircraft energy climbs. The approach, which is based on a nondimensionalization of the equations of motion, is used to evaluatae the appropriateness of forced singular perturbation formulations used in the past for transport and fighter aircraft, and to assess the applicability of energy state approximations and singular perturbation analysis for airbreathing transatmospheric vehicles with hypersonic cruise and orbital capabilities.
The SC "Phobos-Grunt" flight is planned to 2009 in Russia with the purpose to deliver to the Earth the soil samples of the Mars satellite Phobos. The mission will pass under the following scheme [1-4]: the SC flight from the Earth to the Mars, the SC transit on the Mars satellite orbit, the motion round the Mars on the observation orbit and on the quasi-synchronous one [5], landing on Phobos, taking of a ground and start in the direction to the Earth. The implementation of complicated dynamical operations in the Phobos vicinity is foreseen by the project. The SC will be in a disturbance sphere of gravitational fields from the Sun, the Mars and the Phobos. The SC orbit determination is carried out on a totality of trajectory measurements executed from ground tracking stations and measurements of autonomous systems onboard space vehicle relatively the Phobos. As ground measurements the radio engineering measurements of range and range rate are used. There are possible as onboard optical observations of the Phobos by a television system and ranges from the SC up to the Phobos surface by laser locator. As soon as the Phobos orbit accuracy is insufficient for a solution of a problem of landing its orbit determination will be carried out together with determination of the SC orbit. Therefore the algorithms for joint improving of initial conditions of the SC and the Phobos are necessary to determine parameters of the SC relative the Phobos motion within a single dynamical motion model. After putting on the martial satellite orbit, on the Phobos observation orbit, on the quasi-synchronous orbit in the Phobos vicinity the equipment guidance and the following process of the SC orbit determination relatively Phobos requires a priori knowledge of the Phobos orbit parameters with sufficiently high precision. These parameters should be obtained beforehand using both all modern observations and historical ones.
Flight deck-based vision system technologies, such as Synthetic Vision (SV) and Enhanced Flight Vision Systems (EFVS), may serve as a revolutionary crew/vehicle interface enabling technologies to meet the challenges of the Next Generation Air Transportation System Equivalent Visual Operations (EVO) concept - that is, the ability to achieve the safety of current-day Visual Flight Rules (VFR) operations and maintain the operational tempos of VFR irrespective of the weather and visibility conditions. One significant challenge lies in the definition of required equipage on the aircraft and on the airport to enable the EVO concept objective. A motion-base simulator experiment was conducted to evaluate the operational feasibility, pilot workload and pilot acceptability of conducting straight-in instrument approaches with published vertical guidance to landing, touchdown, and rollout to a safe taxi speed in visibility as low as 300 ft runway visual range by use of onboard vision system technologies on a Head-Up Display (HUD) without need or reliance on natural vision. Twelve crews evaluated two methods of combining dual sensor (millimeter wave radar and forward looking infrared) EFVS imagery on pilot-flying and pilot-monitoring HUDs as they made approaches to runways with and without touchdown zone and centerline lights. In addition, the impact of adding SV to the dual sensor EFVS imagery on crew flight performance, workload, and situation awareness during extremely low visibility approach and landing operations was assessed. Results indicate that all EFVS concepts flown resulted in excellent approach path tracking and touchdown performance without any workload penalty. Adding SV imagery to EFVS concepts provided situation awareness improvements but no discernible improvements in flight path maintenance.
The design, building, and testing of the postlanding support systems for a water-landing Assured Crew Return Vehicle (ACRV) are presented. One ACRV will be permanently docked to Space Station Freedom, fulfilling NASA's commitment to Assured Crew Return Capability in the event of an accident or illness. The configuration of the ACRV is based on an Apollo Command Module (ACM) derivative. The 1990-1991 effort concentrated on the design, building, and testing of a one-fifth scale model of the egress and stabilization systems. The objective was to determine the feasibility of (1) stabilizing the ACM out of the range of motions that cause seasickness and (2) the safe and rapid removal of a sick or injured crew member from the ACRV. The development of the ACRV postlanding systems model was performed at the University of Central Florida with guidance from the Kennedy Space Center ACRV program managers. Emphasis was placed on four major areas. First was design and construction of a one-fifth scale model of the ACM derivative to accommodate the egress and stabilization systems for testing. Second was the identification of a water test facility suitable for testing the model in all possible configurations. Third was the construction of the rapid egress mechanism designed in the previous academic year for incorporation into the ACRV model. The fourth area was construction and motion response testing of the attitude ring and underwater parachute systems.
The NASA-funded Pterodactyl project seeks to advance the current state-of-the-art for entry vehicles by developing novel guidance and control technologies for Deployable Entry Vehicles. This paper builds upon the Pterodactyl architecture that employed eight individually articulating flaps with two options for guidance, bank angle modulation with the Fully Numerical Predictor Corrector Entry Guidance (FNPEG) technique and angle of attack and sideslip modulation with FNPEG uncoupled range control. These, with a linear quadratic regulator controller, have previously been presented in separate 3-DOF trajectory simulations, one for translational motion and another for rotational dynamics. This work will show results for integrated 6-DOF simulations, leveraging recent advancements in trajectory simulation tools, namely the Julia-based Genesis package. Results show similar performance for both schemes when compared to the previously presented 3-DOF results once the controllers were adequately tuned. Angle of attack and sideslip modulation necessitated a controller designed at high dynamic pressure conditions, and bank angle modulation necessitated a controller designed at lower dynamic pressure conditions. Notably, angle of attack and sideslip modulation could not achieve the parachute deploy point target Mach number of 2, when only tuning the controller gains, thus future work will be needed to tune guidance specific parameters to achieve acceptable range targeting and guidance command tracking.
In order to develop robust autonomy in spacecraft, it is desirable to develop methods for guiding and controlling arbitrarily-configured spacecraft with any combination of thrusters of various types, i.e. either static reaction control system thrusters and gimbaling thrusters. Scenarios in which autonomous selection of thrusters may be needed include the case of a stuck or inhibited thruster that restricts the motion of the vehicle, vehicles with changing mass properties such as logistics modules or tugs, or vehicles that have imposed constraints on thrusters during docking in order to avoid plume impingement on a space station or adjacent spacecraft. To that end, a spacecraft must be able to rapidly and autonomously reconfigure thruster firing histories and update guidance protocols in accordance with newly imposed constraints. In this paper, a methodology is presented that enables a spacecraft to autonomously select thrusters of any configuration and type in order to optimally match a desired 6-degree-of-freedom navigation and control within the special Euclidean SE(3) framework. The residual motion imposed by off-nominal thruster configurations or thrusters that are not fully controllable is identified by the spacecraft and solved for over time, both for static reaction control systems and for the case of gimbaling thrusters.
The NASA-funded Pterodactyl project seeks to advance the current state-of-the-art for entry vehicles by developing novel guidance and control technologies for Deployable Entry Vehicles. This paper builds upon the Pterodactyl architecture that employed eight individually articulating flaps with two options for guidance, bank angle modulation with the Fully Numerical Predictor Corrector Entry Guidance (FNPEG) technique and angle of attack and sideslip modulation with FNPEG uncoupled range control. These, with a linear quadratic regulator controller, have previously been presented in separate 3-DOF trajectory simulations, one for translational motion and another for rotational dynamics. This work will show results for integrated 6-DOF simulations, leveraging recent advancements in trajectory simulation tools, namely the Julia-based Genesis flight simulation tool. Results show similar performance for both schemes when compared to the previously presented 3-DOF results once the controllers were adequately tuned. Angle of attack and sideslip angle modulation necessitated a controller designed at high dynamic pressure conditions, and bank angle modulation necessitated a controller designed at lower dynamic pressure conditions. Notably, the angle of attack and sideslip modulation approach could not achieve the parachute deploy point target Mach number of 2, when only tuning the controller gains, so future work will be needed to improve guidance and control tracking.
Assembly robots combine the benefits of speed and accuracy with the capability of adaptation to changes in the work environment. However, an impediment to the use of robots is the complexity of the man-machine interface. This interface can be improved by providing a means of using a priori-knowledge and reasoning capabilities for controlling and monitoring the tasks performed by robots. Robots ought to be able to perform complex assembly tasks with the help of only supervisory guidance from human operators. For such supervisory quidance, it is important to express the commands in terms of the effects desired, rather than in terms of the motion the robot must undertake in order to achieve these effects. A suitable knowledge representation can facilitate the conversion of task level descriptions into explicit instructions to the robot. Such a system would use symbolic relationships describing the a priori information about the robot, its environment, and the tasks specified by the operator to generate the commands for the robot.
Fast Fourier Transform (FFT) and Least Square Error (LSE) estimation techniques were applied to the problem of identifying pilot-vehicle dynamic characteristics in flight simulation. A brief investigation of the effects of noise, input bandwidth and system delay upon the FFT and LSE techniques was undertaken using synthetic data. Data from a piloted simulation conducted at NASA Ames Research Center was then analyzed. The simulation was performed in the NASA Ames Research Center Variable Stability CH-47B helicopter operating in fixed-basis simulator mode. The piloting task consisted of maintaining the simulated vehicle over a moving hover pad whose motion was described by a random-appearing sum of sinusoids. The two test subjects used a head-down, color cathode ray tube (CRT) display for guidance and control information. Test configurations differed in the number of axes being controlled by the pilot (longitudinal only versus longitudinal and lateral), and in the presence or absence of an important display indicator called an 'acceleration ball'. A number of different pilot-vehicle transfer functions were measured, and where appropriate, qualitatively compared with theoretical pilot- vehicle models. Some indirect evidence suggesting pursuit behavior on the part of the test subjects is discussed.
Road transportation accounts for 23% of our nation's total energy consumption, 59% of our petroleum consumption and 22% of our total emissions. Electric vehicles (EVs) have much higher well-to-wheel efficiency compared to gasoline vehicles and can reduce our dependence on oil, minimize emissions, improve local air quality and provide a platform for technological innovation and economic growth. While major progress has been made in the development of electric vehicles, their penetration remains under 1 The main hurdles in their widespread adoption are high cost, limited range and long charging times, due to limitations in battery technology. An approach to overcome these hurdles is to substantially reduce the on-board energy storage and instead deliver power wirelessly to the vehicle while it is in motion dynamic wireless power transfer (WPT). For EV charging applications, researchers have traditionally focused on inductive WPT. An issue with inductive WPT systems is that for magnetic flux guidance and shielding, they require ferrite cores, making them expensive, bulky and difficult to embed in the roadway. Under an ARPA-E IDEAS grant, the CU Boulder team has recently developed an alternative approach to WPT based on capacitive coupling that is much smaller, lighter, less expensive, and easier to embed in the roadway than inductive WPT; and have demonstrated a 1.2-kW 6.78-MHz version of this capacitive WPT system on a stationary platform. Under this seed grant program, preliminary work will be carried out on a dynamic capacitive WPT system to prepare a strong proposal for the ARPA-E OPEN 2018 FOA.