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Development of a unified control synthesis methodology for complex and/or non-conventional flight vehicles, and prediction techniques for the handling characteristics of such vehicles are reported. Identification of pilot dynamics and objectives, using time domain and frequency domain methods is proposed.
A conceptual design of the control system for Payload Pointing Systems (PPS) is developed using classic Proportional-Integral-Derivatives (PID) techniques. The major source of system pointing error is due to the disturbance-rich environment of the space station in the form of gimbal baseplate motions. These baseplate vibrations are characterized using Fast Fourier Transform (FFT) techniques. Both time domain and frequency domain dynamic models are developed to assess control system performance. Three basic methods exist for the improvement of PPS pointing performance: increase control system bandwidth, add Image Motion Compensation, and/or reduce (or change) the baseplate disturbance environment.
Some techniques and procedures used in rotor aeroacoustics are examined. The advantages and disadvantages of using the acoustic analogy for noise prediction of helicopter rotors are discussed. The time domain and frequency domain approaches to rotor noise predictions are described. Consideration is given to steady and unsteady thickness noise, loading noise, and quadrupole noise.
The paper presents a method for formulating finite state unsteady aerodynamic models in the time domain from frequency domain unsteady aerodynamics. The method is based on recognizing that the lift deficiency function represents an aerodynamic transfer function and utilizes the Bode plot technique, used in control systems engineering, to construct approximation to the lift deficiency function. Indicial response functions for both fixed wing and rotary wing applications are obtained, using these finite state unsteady aerodynamic models. It is shown that the rotary wing indicial response function is oscillatory and thus it is fundamentally different when compared to the fixed wing indicial response function which is nonoscillatory. Certain aspects of the finite state aerodynamic model are demonstrated by applying it to the flapping dynamics of an articulated helicopter rotor blade. The influence of unsteady aerodynamics on the damping characteristics of the rotor is examined. The same problem is also treated by using a different unsteady aerodynamic model, namely dynamic inflow. Based on a comparison of the results obtained with these two unsteady aerodynamic models, useful conclusions are drawn regarding some fundamental features of these theories.
A summary is presented of the main results obtained during the course of research on output feedback control. The term output feedback is used to denote a controller design approach which does not rely on an observer to estimate the states of the system. Thus, the order of the controller is fixed, and can even be zero order, which amounts to constant gain ouput feedback. The emphasis has been on optimal output feedback. That is, a fixed order controller is designed based on minimizing a suitably chosen quadratic performance index. A number of problem areas that arise in this context have been addressed. These include developing suitable methods for selecting an index of performance, both time domain and frequency domain methods for achieving robustness of the closed loop system, developing canonical forms to achieve a minimal parameterization for the controller, two time scale design formulations for ill-conditioned systems, and the development of convergent numerical algorithms for solving the output feedback problem.
A set of parallel plasma layers is generated by two intersecting microwave pulses in a chamber containing dry air at a pressure comparable to the upper atmosphere. The dependencies of breakdown conditions on the pressure and pulse length are examined. The results are shown to be consistent with the appearance of tail erosion of the microwave pulse caused by air breakdown. A Bragg scattering experiment, using the plasma layers as a Bragg reflector, is then performed. Both time domain and frequency domain measurements of wave scattering are conducted. The experimental results are found to agree very well with the theory.
A set of parallel plasma layers is generated by two intersecting microwave pulses in a chamber containing dry air at a pressure comparable to the upper atmosphere. The dependencies of breakdown conditions on the pressure and pulse length are examined. The results are shown to be consistent with the appearance of tail erosion of microwave pulse caused by air breakdown. A Bragg scattering experiment using the plasma layers as a Bragg reflector is then performed. Both time domain and frequency domain measurements of wave scattering are conducted. The results are found to agree very well with the theory. Moreover, the time domain measurement of wave scattering provides an unambiguous way for determining the temporal evolution of electron density during the first 100 microsec period. A Langmuir double probe is also used to determine the decay rate of electron density during a later time interval (1 to 1.1 ms). The propagation of high power microwave pulses through the air is also studied experimentally. The mechanism responsible for two different degrees of tail erosion were identified. The optimum amplitude of a 1.1 microsec pulse for maximum energy transfer through the air was determined.
Two major tests of the Tethered Satellite System (TSS) engineering and flight units were conducted to demonstrate the functionality of the hardware and software. Deficiencies in the hardware/software integration tests (HSIT) led to a recommendation for more testing to be performed. Selected problem areas of tether dynamics were analyzed, including verification of the severity of skip rope oscillations, verification or comparison runs to explore dynamic phenomena observed in other simulations, and data generation runs to explore the performance of the time domain and frequency domain skip rope observers.
Radar Cross Section (RCS) calculations for flat, perfectly conducting plates are readily available through the use of conventional frequency domain techniques such as the Method of Moments. However, if time domain scattering or wideband frequency domain results are desired, then the Finite Difference Time Domain (FDTD) technique is a suitable choice. In this paper, we present the application of the Finite Difference Time Domain (FDTD) technique to the problem of electromagnetic scattering and RCS calculations from a thin, perfectly conducting plate for a conical cut in the scattering angle phi. RCS calculations versus angle phi will be presented and discussed.
Radar cross section (RCS) calculations for flat, perfectly conducting plates are readily available through the use of conventional frequency domain techniques such as the method of moments. However, if time domain scattering or wideband frequency domain results are desired, then the finite difference time domain (FDTD) technique is a suitable choice. We present the application of the FDTD technique to the problem of electromagnetic scattering and RCS calculations from a thin, perfectly conducting plate for a conical cut in the scattering angle phi. RCS calculations versus angle phi are presented and discussed.
A set of parallel plasma layers is generated by two intersecting microwave pulses in a chamber containing dry air at a pressure comparable to the upper atmosphere. The dependence of the breakdown conditions on the pressure and pulse length is examined. The results are shown to be consistent with the appearance of tail erosion of microwave pulse caused by air breakdown. Bragg scattering experiments, using the plasma layers as a Bragg reflector are then performed. Both time domain and frequency domain measurements of wave scattering are conducted. The experimental results are found to agree very well with the theory. Moreover, the time domain measurement of wave scattering provides an unambiguous way for determining the temporal evolution of electron density during the first 100-microsec period.
An iterative procedure is proposed for computing the commanded state trajectories and controls that guide a possibly multiaxis, time-varying, nonlinear system with transmission zero dynamics through a given arbitrary sequence of control points. The procedure is initialized by the system inverse with the transmission zero effects nulled out. Then the 'steady state' solution of the perturbation model with the transmission zero dynamics intact is computed and used to correct the initial zero-free solution. Both time domain and frequency domain methods are presented for computing the steady state solutions of the possibly nonminimum phase transmission zero dynamics. The procedure is illustrated by means of linear and nonlinear examples.
This guide describes the input data required, for steady or unsteady aerodynamic and aeroelastic analysis of propellers and the output files generated, in using PROP3D. The aerodynamic forces are obtained by solving three dimensional unsteady, compressible Euler equations. A normal mode structural analysis is used to obtain the aeroelastic equations, which are solved using either time domain or frequency domain solution method. Sample input and output files are included in this guide for steady aerodynamic analysis of single and counter-rotation propellers, and aeroelastic analysis of single-rotation propeller.
Measurement of the microgravity acceleration environment on the International Space Station will be accomplished by two accelerometer systems. The Microgravity Acceleration Measurement System will record the quasi-steady microgravity environment, including the influences of aerodynamic drag, vehicle rotation, and venting effects. Measurement of the vibratory/transient regime comprised of vehicle, crew, and equipment disturbances will be accomplished by the Space Acceleration Measurement System-II. Due to the dynamic nature of the microgravity environment and its potential to influence sensitive experiments, Principal Investigators require distribution of microgravity acceleration in a timely and straightforward fashion. In addition to this timely distribution of the data, long term access to International Space Station microgravity environment acceleration data is required. The NASA Glenn Research Center's Principal Investigator Microgravity Services project will provide the means for real-time and post experiment distribution of microgravity acceleration data to microgravity science Principal Investigators. Real-time distribution of microgravity environment acceleration data will be accomplished via the World Wide Web. Data packets from the Microgravity Acceleration Measurement System and the Space Acceleration Measurement System-II will be routed from onboard the International Space Station to the NASA Glenn Research Center's Telescience Support Center. Principal Investigator Microgravity Services' ground support equipment located at the Telescience Support Center will be capable of generating a standard suite of acceleration data displays, including various time domain and frequency domain options. These data displays will be updated in real-time and will periodically update images available via the Principal Investigator Microgravity Services web page.
The Hubble Space Telescope (HST) normally requires three gyroscopes for three-axis rate control. The loss of the Space Shuttle Columbia on STS-107 resulted in the cancellation of a shuttle-based HST Servicing Mission 4. Therefore, HST must operate using the on-board hardware until an alternate means of servicing can be accomplished. The probability of gyro failure indicates that fewer than three gyros will be operable before any servicing mission can be performe& To mitigate this, and to extend the HST life expectancy, a rate estimation and control algorithm was developed that requires two gyros to measure rate about two axes, with the remaining axis rate estimated using one of three alternate sensors. Three-axis magnetometers (MSS) are used for coarse rate estimation during large maneuvers and during occultations of other sensors. Fixed-Head Star Trackers (FHSTs) are used for rate estimation during safe mode recovery and during transition to science operations. Fine rate estimation during science operations is performed using the Fine Guidance Sensors (FGSs). The FHST mode (T2G) relies on star vectors as measured by the FHSTs to estimate vehicle rate about the axis not measured by the gyros. Since the FHSTs were not designed to estimate body rate, this method involves a unique set of problems that had to be overcome in the final design, such as the effect of FHST break tracks and moving targets on rate estimation. The solutions to these problems, as well as a detailed description of the design and implementation of the rate estimation are presented Also included are the time domain and frequency domain analysis of the T2G control law. A high fidelity HST simulator (HSTSIM) was used to verify T2G performance prior to on-orbit use. Results of these simulations are also presented. Finally, analysis of actual T2G on-orbit test results is presented for design validation.
The equations for frequency-domain multiple scattering are derived for a scalar or electromagnetic plane wave incident on a collection of particles at known positions, and in the time-domain for a plane wave pulse incident on the same collection of particles. The calculation is carried out for five different combinations of wave types and particle types of increasing geometrical complexity. The results are used to illustrate and discuss a number of physical and mathematical characteristics of multiple scattering in the frequency- and time-domains. We argue that frequency-domain multiple scattering is a purely mathematical construct since there is no temporal sequencing information in the frequency-domain equations and since the multi-particle path information can be dispelled by writing the equations in another mathematical form. However, multiple scattering becomes a definite physical phenomenon in the time-domain when the collection of particles is illuminated by an appropriately short localized pulse.
The Frequency Domain Input Synthesis procedure is used in identifying the stability and control derivatives of an aircraft. By using a frequency-domain approach, one can handle criteria that are not easily handled by the time-domain approaches. Numerical results are presented for optimal elevator deflections to estimate the longitudinal stability and control derivatives subject to root-mean square constraints on the input. The applicability of the steady state optimal inputs to finite duration flight testing is investigated. The steady state approximation of frequency-domain synthesis is good for data lengths greater than two time cycles for the short period mode of the aircraft longitudinal motions. Phase relationships between different frequency components become important for shorter data lengths. The frequency domain inputs are shown to be much better than the conventional doublet inputs.
The baseline stability margins for NASA's Space Launch System (SLS) launch vehicle were generated via the classical approach of linearizing the system equations of motion and determining the gain and phase margins from the resulting frequency domain model. To improve the fidelity of the classical methods, the linear frequency domain approach can be extended by replacing static, memoryless nonlinearities with describing functions. This technique, however, does not address the time varying nature of the dynamics of a launch vehicle in flight. An alternative technique for the evaluation of the stability of the nonlinear launch vehicle dynamics along its trajectory is to incrementally adjust the gain and/or time delay in the time domain simulation until the system exhibits unstable behavior. This technique has the added benefit of providing a direct comparison between the time domain and frequency domain tools in support of simulation validation.