Randolph H. Cabell

Experimental Feedback Control of Flow Induced Cavity Tones

Discrete-time, linear quadratic methods were used to design feedback controllers for reducing tones generated by flow over a cavity. The dynamics of a synthetic-jet actuator mounted at the leading edge of the cavity as observed by two microphones in the cavity were modeled over a broad frequency range using state space models computed from experimental data. Variations in closed loop performance as a function of model order, control order, control bandwidth, and state estimator design were studied using a cavity in the Probe Calibration Tunnel at NASA Langley Research Center. The controller successfully reduced the levels of multiple cavity tones at the tested flow speeds of Mach 0.275,0.35, and 0.45. In some cases, the closed loop results were limited by excitation of sidebands of the cavity tones, or the creation of new tones at frequencies away from the cavity tones. The models were not able to account for nonlinear dynamics, such as interactions between tones at different frequencies. Nonetheless, the results validate the combination of optimal control and experimentally generated state space models for the cavity tone problem.

Development of an Experimental Rig for Investigation of Higher Order Modes in Ducts

Continued progress to reduce fan noise emission from high bypass ratio engine ducts in aircraft increasingly relies on accurate description of the sound propagation in the duct. A project has been undertaken at NASA Langley Research Center to investigate the propagation of higher order modes in ducts with flow. This is a two-pronged approach, including development of analytic models (the subject of a separate paper) and installation of a laboratory-quality test rig. The purposes of the rig are to validate the analytical models and to evaluate novel duct acoustic liner concepts, both passive and active. The dimensions of the experimental rig test section scale to between 25% and 50% of the aft bypass ducts of most modern engines. The duct is of rectangular cross section so as to provide flexibility to design and fabricate test duct liner samples. The test section can accommodate flow paths that are straight through or offset from inlet to discharge, the latter design allowing investigation of the effect of curvature on sound propagation and duct liner performance. The maximum air flow rate through the duct is Mach 0.3. Sound in the duct is generated by an array of 16 high-intensity acoustic drivers. The signals to the loudspeaker array are generated by a multi-input/multi-output feedforward control system that has been developed for this project. The sound is sampled by arrays of flush-mounted microphones and a modal decomposition is performed at the frequency of sound generation. The data acquisition system consists of two arrays of flush-mounted microphones, one upstream of the test section and one downstream. The data are used to determine parameters such as the overall insertion loss of the test section treatment as well as the effect of the treatment on a modal basis such as mode scattering. The methodology used for modal decomposition is described, as is a description of the mode generation control system. Data are presented which demonstrate the performance of the controller to generate the desired mode while suppressing all other cut on modes in the duct.

Design and Control of a Morphing Chevron for Takeoff and Cruise Noise Reduction

‡Commercial high-bypass ratio turbofan engines generate high levels of noise as the jet exhaust mixes with the ambient air. Serrated aerodynamic devices, known as chevrons, positioned along the trailing edges of a jet engine’s primary and secondary exhaust nozzle have been shown to reduce jet noise at take off as well as shock-cell noise at cruise. Their optimum shape is a finely tuned compromise between noise-benefit and thrust-loss. We report here the development of a powered variable geometry chevron (PVGC) capable of transitioning between take off and cruise shapes. Actuators composed of thermally active nickel-titanium (NiTinol) shape memory alloy drive the shape change. A full scale PVGC was tested under representative flow conditions in Boeing’s Nozzle Test Facility (NTF). In this test a simple proportional-integral control system provided continuous control of the VGC tip immersion between take off and cruise conditions. The collected data supports the aerodynamic, thermal, and mechanical design of the PVGC as well as the control system approach.

A principal component feedforward algorithm for active noise control: flight test results

An in-flight evaluation of a principal component algorithm for feedforward active noise control is discussed. Cabin noise at the first three harmonics of the blade passage frequency (103 Hz) of a Raytheon-Beech 1900D twin turboprop aircraft was controlled using 21 pairs of inertial force actuators bolted to the ring frames of the aircraft; 32 microphones provided error feedback inside the aircraft cabin. In a single frequency noise control test, the blade passage frequency was reduced by 15 dB averaged across the microphone array. When controlling the first three harmonics simultaneously, reductions of 11 dB at 103 Hz, 1.5 dB at 206 Hz, and 2.8 dB at 309 Hz were obtained. For single frequency feedforward control problems, the principal component algorithm is shown to be useful for reducing the computational burden and simplifying the implementation of control effort penalties in high channel count control systems. Good agreement was found between the in-flight behavior of the controller and the predicted optimal control solution.

Artificial neural networks for predicting nonlinear dynamic helicopter loads

This paper proposes using an artificial neural network (ANN) to predict the loads in critical components based on flight variable information that can be easily measured. The artificial neural network learns the relationship between flight variables and component loads through exposure to a database of flight variable records and corresponding load histories taken from an instrumented military helicopter undergoing standard maneuvers. Eight standard flight variables are used as inputs for predicting the time varying mean and oscillatory components of the tailboom bending load and the pitch link load for seven flight maneuvers

Identification of Helicopter Noise Using a Neural Network

An artificial neural network (ANN) has been trained to distinguish between the noise of two helicopters. The performance of the ANN is compared with that of a conventional recognition system. The conventional system uses the ratio of the main-rotor blade passage frequency (bpf) to the tail-rotor bpf. The ANN was trained to use similar main/tail-rotor information, in addition to information describing the distribution of spectral peaks of the main rotor. It is shown that this additional information allows the ANN to distinguish between the helicopters when tail-rotor noise is removed from the spectrum. The performance of the two methods is given as a function of signal-to-noi se strength, and propagation distance, using a model of atmospheric sound propagation. The conventional method outperforms the ANN when main- and tail-rotor noise are present, but the conventional method cannot identify helicopters when tail-rotor noise is removed. At 20-dB signal-to-noise ratio (SNR), when tail-rotor noise is not present in the spectrum, the ANN correctly identifies the helicopters 100% of the time, compared to 50% for the conventional method. The performance of the ANN drops as signal strength decreases. At 8-dB SNR, the ANN is correct 77% of the time, while at 0 dB it is correct 58% of the time. Similar results are obtained for the performance when the signal is propagated through the model of the atmosphere.

Controller Complexity for Active Control of Turbulent Boundary-Layer Noise from Panels

An experimental study of feedback controller complexity vs noise reduction performance for active structural acoustic control of turbulent-boundary-layer (TBL)-induced sound radiation from a panel is described. The reduction of total radiated sound power as a function of the number of actuators, sensors, and controller cost function on a mock aircraft sidewall subjected to TBL excitation are discussed. The results demonstrate total radiated sound power reductions of 15 dB at resonances and 10 dB integrated over 150-1000 Hz for a three-actuator and 15-sensor case. The controller configuration was then simplified to one actuator and four sensors (summed outputs) and was found to produce 10-15-dB reductions in sound power at resonances and 9 dB integrated over the control bandwidth. This result demonstrates the potential for achieving significant reductions in radiated sound power with a relatively simple actuator/sensor topology.

Structural acoustic control of plates with variable boundary conditions: Design methodology

A method for optimizing a structural acoustic control system subject to variations in plate boundary conditions is provided. The assumed modes method is used to build a plate model with varying levels of rotational boundary stiffness to simulate the dynamics of a plate with uncertain edge conditions. A transducer placement scoring process, involving Hankel singular values, is combined with a genetic optimization routine to find spatial locations robust to boundary condition variation. Predicted frequency response characteristics are examined, and theoretically optimized results are discussed in relation to the range of boundary conditions investigated. Modeled results indicate that it is possible to minimize the impact of uncertain boundary conditions in active structural acoustic control by optimizing the placement of transducers with respect to those uncertainties.

Active Control of Turbulent Boundary Layer Induced Sound Radiation from Multiple Aircraft Panels

The objective of this work is to experimentally investigate active structural acoustic control of turbulent boundary layer (TBL) induced sound radiation from multiple panels on an aircraft sidewall. One possible approach for controlling sound radiation from multiple panels is a multi-input/multi-output scheme which considers dynamic coupling between the panels. Unfortunately, this is difficult for more than a few panels, and is impractical for a typical aircraft which contains several hundred such panels. An alternative is to implement a large number of independent control systems. Results from the current work demonstrate the feasibility of reducing broadband radiation from multiple panels utilizing a single-input/single-output (SISO) controller per bay, and is the first known demonstration of active control of TBL induced sound radiation on more than two bays simultaneously. The paper compares sound reduction for fully coupled control of six panels versus independent control on each panel. An online adaptive control scheme for independent control is also demonstrated. This scheme will adjust for slow time varying dynamic systems such as fuselage response changes due to aircraft pressurization, etc.

Real-Time Adaptive Control of Flow-Induced Cavity Tones (Invited)

An adaptive generalized predictive control (GPC) algorithm was formulated and applied to the cavity flow-tone problem. The algorithm employs gradient descent to update the GPC coefficients at each time step. The adaptive control algorithm demonstrated multiple Rossiter mode suppression at fixed Mach numbers ranging from 0.275 to 0.38. The algorithm was also able to maintain suppression of multiple cavity tones as the freestream Mach number was varied over a modest range (0.275 to 0.29). Controller performance was evaluated with a measure of output disturbance rejection and an input sensitivity transfer function. The results suggest that disturbances entering the cavity flow are colocated with the control input at the cavity leading edge. In that case, only tonal components of the cavity wall-pressure fluctuations can be suppressed and arbitrary broadband pressure reduction is not possible. In the control-algorithm development, the cavity dynamics are treated as linear and time invariant (LTI) for a fixed Mach number. The experimental results lend support this treatment.