Walter F. O'Brien

Active control of aircraft engine inlet noise using compact sound sources and distributed error sensors

An active noise control system using a compact sound source is effective to reduce aircraft engine duct noise. The fan noise from a turbofan engine is controlled using an adaptive filtered-x LMS algorithm. Single multi channel control systems are used to control the fan blade passage frequency (BPF) tone and the BPF tone and the first harmonic of the BPF tone for a plane wave excitation. A multi channel control system is used to control any spinning mode. The multi channel control system to control both fan tones and a high pressure compressor BPF tone simultaneously. In order to make active control of turbofan inlet noise a viable technology, a compact sound source is employed to generate the control field. This control field sound source consists of an array of identical thin, cylindrically curved panels with an inner radius of curvature corresponding to that of the engine inlet. These panels are flush mounted inside the inlet duct and sealed on all edges to prevent leakage around the panel and to minimize the aerodynamic losses created by the addition of the panels. Each panel is driven by one or more piezoelectric force transducers mounted on the surface of the panel. The response of the panel to excitation is maximized when it is driven at its resonance; therefore, the panel is designed such that its fundamental frequency is near the tone to be canceled, typically 2000-4000 Hz.

Active control of fan noise from a turbofan engine

A three-channel active control system is applied to an operational turbofan engine to reduce tonal noise produced by both the fan and the high-pressure compressor. The control approach is the feedforward filtered-x least-mean-square algorithm implemented on a digital signal processing board. Reference transducers mounted on the engine case provide blade passing and harmonics frequency information to the controller. Error information is provided by large area microphones placed in the acoustic far field. To minimize the error signal, the controller actuates loudspeakers mounted on the inlet to produce destructive interference

Dual-Mode Combustion Experiments with an Integrated Aeroramp-Injector/Plasma-Torch Igniter

Results from combustion experiments in a direct-connect supersonic combustor facility are presented. Successful ignition and sustained combustion of both hydrogen and ethylene fuels were achieved using an integrated aerorampinjector/plasma-torch igniter configuration. A Mach 2 nozzle was used to obtain flow simulating Mach ≈ 4 flight conditions at 27 km, at a total temperature of 1000 K and a static pressure of 42 kPa. Combustion was achieved at (global) equivalence ratios between 0.08 and 0.31 for hydrogen and 0.13 and 0.47 for ethylene, with corresponding maximum combustor pressure rises of about a factor of 4.0. One-dimensional performance analysis of the test data indicates combustion efficiencies as high as 75% for both fuels, in the leanest conditions tested. Off-design flight conditions were tested by varying the freestream air total temperature. Supersonic combustion was achieved at total temperatures as low as 530 K with hydrogen and 680 K with ethylene.

Preliminary Design for Embedded Engine Systems

This paper presents a collaborative research program to advance the analysis and design capabilities for embedded engines in Blended Wing Body Boundary Layer Ingesting (BWB- BLI aircraft. Embedded engine propulsion systems for hybrid wing/body aircraft face several key technical issues stemming from the ingestion of the low-momentum boundary layer to the close coupling of the airframe and propulsion systems in ways that challenge current design procedures and require a system design perspective. To address these challenges, a new high performance flow-control-enabled embedded engine inlet was designed. The inlet bleed flow control actuation was integrated with the aircraft Environmental Control System (ECS), thus providing a viable robust inlet design with improved system performance while supplying the flow required by the ECS. Without flow control, a thick ingested boundary layer leads to high levels of distortion at the AIP plane and impacts the overall fan performance and operability. The impact of inlet distortion on fan performance was estimated through numerical simulations based on unsteady flow simulations. Of particular interest, is the increased flow unsteadiness in the fan passages and its associated detrimental impact on blade aeromechanics and acoustic responses. The proposed flow control solution is predicted to have significant reduction in the harmonic content associated with blade strain. In addition, reductions in total radiated acoustic power are predicted to be achievable, especially at well cut-on modes, significantly reducing aircraft noise at approach conditions. These results suggest that inlet flow control technology has the potential to address the challenges of BWB-BLI aircraft.

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.

Stereoscopic PIV Measurements of Swirl Distortion on a Full-Scale Turbofan Engine Inlet

There is a present need for simulating and measuring the inlet swirl distortion generated by airframe/engine system interactions to identify potential degradation in fan performance and operability in a full-scale, ground testing environment. Efforts are described to address this need by developing and characterizing methods for complex, prescribed distortion patterns. A relevent inlet swirl distortion profile was generated by a novel new method, dubbed the StreamVane method, and measured in a small scale tunnel using stereoscopic particle image velocimetry (PIV) as a precursor for swirl distortion generation and characterization in an operating turbofan research engine. The StreamVane prototype tested was made of ABS plastic using additive manufacturing and was used to generate swirl distortion patterns that mimick boundary layer ingesting engine inlets with a pair of tightly wound vortices and swirl angle magnitudes up to 15°. Diagnostic development efforts for distortion measurements within the research engine paralleled the StreamVane characterization. The system used for research engine PIV measurements is described along with data obtained in the wake of a total pressure distortion screen for engine conditions at idle and 80% corrected fan speed engine power settings. Data reduction algorithms are put forth to reduce spurious velocity vectors and uncertainty estimations specific to the inlet distortion test rig are made. Results indicate that the methods developed may be used to both generate and characterize complex distortion profiles at the aerodynamic interface plane, providing new information about airframe/engine integration.

Integrated Aeroramp-Injector/Plasma-Torch Igniter for Methane and Ethylene Fueled Scramjets

An integrated aeroramp-injector/plasma-torch igniter configuration has been experimentally tested in conditions simulating scramjet flight at a Mach number of 4. Ethylene (C2H4) and methane (CH4) were used as fuels with air as the torch feedstock gas. A range of equivalence ratios (φ) were tested to determine operability limits. Successful combustion was determined by observing wall-static pressure rise in the facility, as well as wall temperature increases. A long isolator in the facility allowed for dual-mode operation, and both subsonic and supersonic combustion modes with good efficiency were observed with ethylene fuel. Poor results were obtained with methane in the same configuration as was successful with ethylene. The GASP numerical solver was used to evaluate the suitability of the plasma torch location with respect to the aeroramp injector center. An analysis of the mixing of each fuel through the aeroramp indicated that the selected torch location for methane was not ideal, as opposed to the ethylene case, for which it was predicted as good. The predictions were in accord with the experimental results. This lends confidence to the idea that CFD can be used to help select a better configuration for methane.

Active Control of Flow in Serpentine Inlets for Blended Wing-Body Aircraft

The Blended Wing-Body commercial aircraft concept promises performance improvements in noise, emissions, and fuel consumption. Highly integrated airframepropulsion systems featuring embedded engines offer further improvements. Embedded engine systems are envisioned which require Boundary Layer Ingesting (BLI) serpentine inlets to provide the needed airflow to the engine. Due to the ingestion of a large boundary layer as well as the geometry of the serpentine inlet, significant flow distortions are developed that will affect engine performance and the stability of the fan. A bleed flow control system was tested that utilized no more than 2% of the total inlet flow. Two bleed slots were employed, one near the entrance of the BLI inlet and one near its aft exit. The bleed system successfully reduced inlet distortions by as much as 30%, implying improvements in stall margin and engine performance.

Scramjet Operability Range Studies of a Multifuel Integrated Aeroramp Injector/Plasma Igniter

The integrated aerodynamic ramp injector plasma torch igniter was tested in a M∞ = 2, un-vitiated, heated flow facility as a combustion system. The facility operated at a total temperature of 1000 K and total pressure of 3.25 atm. Hydrogen (H2) and ethylene (C2H4) were used as fuels and a wide range of global equivalence ratios were tested. The main data obtained were wall static pressure measurements. Wall temperature measurements and visual observations of combustion were also acquired. Ethylene operability limits were determined to be 0.14 < φ < 0.48. The hydrogen fuel data for the aeroramp/torch system was compared to data from a physical 10o unswept ramp injector. Comparable performance was obtained with the two arrangements. With hydrogen as the fuel and the aeroramp/torch system, the effect of varying the tunnel total temperature was investigated. Good combustion was achieved with air total temperatures as low as 530K. The pressure profiles were analyzed using the Ramjet Propulsion Analysis (RJPA) code. Results indicate that both supersonic and dual-mode ramjet combustion were achieved. Combustion efficiencies varied with φ from a high of about 75% to a low of about 45% at the highest φ. With a theoretical diffuser and nozzle assumed for the configuration and engine, thrust was computed for each fuel. Air specific impulse varied from a low of about 9 to a high of about 24 seconds for the To = 1000K test condition. The GASP code was used to numerically simulate the injection and mixing process of the two fuels and the results were used to evaluate the suitability of the selected plasma torch locations.