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Integrated system of stand-along computer programs, DYLOFLEX, analyzes dynamic loads on flexible aircraft with active controls. DYLOFLEX capabilities include calculating dynamic loads due to continuous atmospheric turbulence, discrete gusts, and discrete control inputs. Each of the eight individual DYLOFLEX programs may be used alone or in conjunction with other DYLOFLEX programs.
Conceptual and wind-tunnel programs leading to the development of technology for applying active controls to the suppression of flutter indicate that these methods may prove to be a powerful tool in providing required safety margins for flutter in future high-performance supersonic aircraft. The nature of flutter considerations in the design of the United States SST prototype aircraft has been described as an example of the type of application where active flutter suppression shows promise. Although this new technology is emerging, several years of additional development will be required in order to bring the technology to a complete state of readiness, particularly for civil applications.
Related control-configured vehicle design and system considerations are considered. Representative applications of active control for fighter and transport aircraft are included.
An active noise control model has been evaluated for reducing aircraft interior noise. The structural noise transmission properties of an aircraft fuselage were modelled as a flexible cylinder excited by external acoustic dipoles simulating the noise produced by twin propellers. The amplitudes of an internal distribution of monopole control sources were determined such that the area-weighted mean square acoustic pressure was minimized in the propeller plane. The noise control model was evaluated at low frequencies corresponding to the blade passage frequency and first few harmonics of a typical turbo-prop aircraft. Interior noise reductions of 20 25 dB were achieved, over a substantial region of the cylindrical cross-section, with just a few monopole control sources. The most favorable interior noise reductions were achieved when the active noise control model was used in combination with propeller source phasing.
Active control of sound radiation from a rectangular panel by two different methods has been experimentally studied and compared. In the first method a single control force applied directly to the structure is used with a single error microphone located in the radiated acoustic field. Global attenuation of radiated sound was observed to occur by two main mechanisms. For 'on-resonance' excitation, the control force had the effect of increasing the total panel input impedance presented to the nosie source, thus reducing all radiated sound. For 'off-resonance' excitation, the control force tends not significantly to modify the panel total response amplitude but rather to restructure the relative phases of the modes leading to a more complex vibration pattern and a decrease in radiation efficiency. For acoustic control, the second method, the number of acoustic sources required for global reduction was seen to increase with panel modal order. The mechanism in this case was that the acoustic sources tended to create an inverse pressure distribution at the panel surface and thus 'unload' the panel by reducing the panel radiation impedance. In general, control by structural inputs appears more effective than control by acoustic sources for structurally radiated noise.
Active flow control in the form of periodic zero-mass-flux excitation was applied at several regions on the leading edge and trailing edge flaps of a simplified high-lift system t o delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approx.= 10) and low frequency amplitude modulation (F(+)AM approx.= 1) of the high frequency excitation were used for control. Preliminary efforts to combine leading and trailing edge flap excitations are also reported.
A computer program, ISAC, is described which calculates the stability and response of a flexible airplane equipped with active controls. The equations of motion relative to a fixed inertial coordinate system are formulated in terms of the airplane's rigid body motion and its unrestrained normal vibration modes. Unsteady aerodynamic forces are derived from a doublet lattice lifting surface theory. The theoretical basis for the program is briefly explained together with a description of input data and output results.
The transverse vibrations of beams is discussed and a methodology for the design of an active damping device is given. The Bernoulli-Euler equation is used to derive a transcendental transfer function, which relates a torque applied at one end of the beam to the rotational position and velocity at that point. The active damping device consists of a wire, a linear actuator and a short torque arm attached to one end of the beam. The action of the actuator varies a tension in the wire and creates a torque which opposes the rotation of the beam and thus damps vibration. A design procedure for such an active damper is given. This procedure shows the relationships and trade-offs between the actuator stroke, power required, stress levels in the wire and beam and the geometry of the beam and wire. It is shown that by consideration of the frequency response at the beam natural frequencies, the aforementioned relationships can be greatly simplified. Similarly, a simple way of estimating the effective damping ratios and eigenvalue locations of actively controlled beams is presented.
The integrated application of active controls (IAAC) technology to an advanced subsonic transport is reported. Supplementary technical data on the following topics are included: (1) 1990's avionics technology assessment; (2) function criticality assessment; (3) flight deck system for total control and functional features list; (4) criticality and reliability assessment of units; (5) crew procedural function task analysis; and (6) recommendations for simulation mechanization.
The primary objective of the Test Active Control Technology (ACT) System laboratory tests was to verify and validate the system concept, hardware, and software. The initial lab tests were open loop hardware tests of the Test ACT System as designed and built. During the course of the testing, minor problems were uncovered and corrected. Major software tests were run. The initial software testing was also open loop. These tests examined pitch control laws, wing load alleviation, signal selection/fault detection (SSFD), and output management. The Test ACT System was modified to interface with the direct drive valve (DDV) modules. The initial testing identified problem areas with DDV nonlinearities, valve friction induced limit cycling, DDV control loop instability, and channel command mismatch. The other DDV issue investigated was the ability to detect and isolate failures. Some simple schemes for failure detection were tested but were not completely satisfactory. The Test ACT System architecture continues to appear promising for ACT/FBW applications in systems that must be immune to worst case generic digital faults, and be able to tolerate two sequential nongeneric faults with no reduction in performance. The challenge in such an implementation would be to keep the analog element sufficiently simple to achieve the necessary reliability.
Both classical and state-space synthesis methods for active control of flexible space structures in the presence of persistent disturbances are presented. The methods exploit the so-called internal model principle for asymptotic disturbance rejection. A generic example of flexible space structures is used to illustrate the simplicity of the proposed design methodologies. The concept of a disturbance rejection filter dipole is introduced from a classical control viewpoint. It is shown that the proposed design methods will invariably make use of non-minimum-phase compensation for a class of noncolocated control problems. The need for tradeoffs between performance and parameter robustness is discussed.
A very general formulation has been developed for the analysis of complex flexible-aircraft configurations in presence of active control. The formulation is incorporated into a computer program, Flight Control Analysis Program (FCAP), which is designed in a modular fashion to incorporate aircraft dynamics and aerodynamics for complex configurations, as well as sensor, actuator and control logic element dynamics. Formulation of the total aircraft dynamic system is accomplished in matrix form by casting equations in state vector format. The system stability and performance are determined in either the frequency or time domain using classical analysis techniques. Fully unsteady aerodynamics is used for the evaluation of the flutter characteristics as well as the gust response of the aircraft.
Relaxed static stability and stability augmentation with active controls were investigated for subsonic transport aircraft. Analytical and simulator evaluations were done using a contemporary wide body transport as a baseline. Criteria for augmentation system performance and unaugmented flying qualities were evaluated. Augmentation control laws were defined based on selected frequency response and time history criteria. Flying qualities evaluations were conducted by pilots using a moving base simulator with a transport cab. Static margin and air turbulence intensity were varied in test with and without augmentation. Suitability of a simple pitch control law was verified at neutral static margin in cruise and landing flight tasks. Neutral stability was found to be marginally acceptable in heavy turbulence in both cruise and landing conditions.
The final design and hardware fabrication was completed for an active control system capable of the required flutter suppression, compatible with and ready for installation in the NASA aeroelastic research wing number 1 (ARW-1) on Firebee II drone flight test vehicle. The flutter suppression system uses vertical acceleration at win buttock line 1.930 (76), with fuselage vertical and roll accelerations subtracted out, to drive wing outboard aileron control surfaces through appropriate symmetric and antisymmetric shaping filters. The goal of providing an increase of 20 percent above the unaugmented vehicle flutter velocity but below the maximum operating condition at Mach 0.98 is exceeded by the final flutter suppression system. Results indicate that the flutter suppression system mechanical and electronic components are ready for installation on the DAST ARW-1 wing and BQM-34E/F drone fuselage.
Because of inherent low damping and high flexibility, large space structures may require some form of active control of their dynamics. Because of the apparent inability to accurately model the dynamics of these structures, methods for parameter adaptive control are now being developed at Langley. This approach uses a digital computer to process discrete sensor data, identify modal parameters, calculate modal control gains, and then convert the modal forces to real forces. Some of the problems considered are: (1) the possibility that there may be modes to control with limited amounts of hardware, and (2) the required accuracy of identified structural parameters.
Work done under the Energy Efficient Transport project in the field of advanced aerodynamics and active controls is summarized. The project task selections focused on the following: the investigation of long-duct nacelle shape variation on interference drag; the investigation of the adequacy of a simple control law for the elastic modes of a wing; the development of the aerodynamic technology at cruise and low speed of high-aspect-ratio supercritical wings of high performance; and the development of winglets for a second-generation jet transport. All the tasks involved analysis and substantial wind tunnel testing. The winglet program also included flight evaluation. It is considered that the technology base has been built for the application of high-aspect-ratio supercritical wings and for the use of winglets on second-generation transports.
The development of an advanced pitch active control system (PACS) and a reduced area horizontal tail for a wide-body jet transport (L-1011) with a flying horizontal stabilizer is discussed. The advanced PACS control law design objectives were to provide satisfactory handling qualities for aft c.g. flight conditions to negative static stability margins of 10 percent and to provide good maneuver control column force gradients for nonlinear stability flight conditions. Validity of the control laws were demonstrated by piloted flight simulation tests on the NASA Langley Visual Motion Simulator. Satisfactory handling qualities were actually demonstrated to a negative 20 percent static stability margin. The PACS control laws were mechanized to provide the system architecture that would be suitable for an L-1011 flight test program to a negative stability margin of 3 percent which represents the aft c.g. limits of the aircraft. Reduced area horizontal tail designs of 30 and 38 percent with respect to the L-1011 standard tail were designed, fabricated and wind tunnel tested. Drag reductions and weight savings of the 30 percent smaller tail would provide an L/D benefit of about 2% and the 38% small tail L/D benefit would be about 3 percent. However, forward c.g. limitations would have to be imposed on the aircraft because the maximum horizontal tail lift goal was not achieved and sufficient aircraft nose-up control authority was not available. This limitation would not be required for a properly designed new aircraft.
An account is given of an effort to reduce helicopter transmission noise by 10 dB, using active methods, as part of the NASA-Lewis/U.S. Army Propulsion Directorate Advanced Rotorcraft Transmission technology integration and demonstration program. The transmission used as a test stand is that of the CH-47C forward rotor. Attention is presently given to the active control system's actuators, sensors, and control algorithms.