Trailing Edge Flow Field Research Articles

Mitigation of Turbine Vane Shock Waves Through Trailing Edge Cooling

The immense market demand on the high efficiency and lightweight aero-engines results in designs with compact high-pressure turbine stages experiencing supersonic flow field. In supersonic turbines, shocks appear at the vane trailing edges. The interaction of these shock with the neighbouring airfoils and blades on the adjacent rotor row and consequently create considerable amount of losses on the aerodynamic performance of the turbine. Moreover, periodic excitation created by the interaction of the shock waves and the motion of the turbine rotor causes fatigue problems and reduces the lifetime of the engine. Current study aims to alter the vane shock waves through blowing at the trailing edge. In order to characterize the effect of active blowing on the trailing edge flow field, a series of URANS simulations were conducted on OpenFOAM solver platform. Various blowing schemes were simulated over a simplified trailing edge geometry exposed in supersonic flow. The computations were compared in terms of shock intensity, oscillation frequency and exerted pressure forcing over the downstream components. The results showed that unsteady trailing edge blowing were able to modify the fluctuations observed on the shocks by altering the shock intensity, angle and frequency of oscillations. The classification of the wake unsteadiness, i.e. vortex shedding, in terms of trailing edge characteristics were also accomplished through frequency domain analysis of simulations.

Numerical and experimental investigation of a beveled trailing-edge flow field and noise emission

Efficient tools and methodology for the prediction of trailing-edge noise experience substantial interest within the wind turbine industry. In recent years, the Lattice Boltzmann Method has received increased attention for providing such an efficient alternative for the numerical solution of complex flow problems. Based on the fully explicit, transient, compressible solution of the Lattice Boltzmann Equation in combination with a Ffowcs-Williams and Hawking aeroacoustic analogy, an estimation of the acoustic radiation in the far field is obtained. To validate this methodology for the prediction of trailing-edge noise, the flow around a flat plate with an asymmetric 25° beveled trailing edge and obtuse corner in a low Mach number flow is analyzed. Flow field dynamics are compared to data obtained experimentally from Particle Image Velocimetry and Hot Wire Anemometry, and compare favorably in terms of mean velocity field and turbulent fluctuations. Moreover, the characteristics of the unsteady surface pressure, which are closely related to the acoustic emission, show good agreement between simulation and experiment. Finally, the prediction of the radiated sound is compared to the results obtained from acoustic phased array measurements in combination with a beamforming methodology. Vortex shedding results in a strong narrowband component centered at a constant Strouhal number in the acoustic spectrum. At higher frequency, a good agreement between simulation and experiment for the broadband noise component is obtained and a typical cardioid-like directivity is recovered.

High-lift airfoil trailing edge separation control using a single dielectric barrier discharge plasma actuator

Control of flow separation from the deflected flap of a high-lift airfoil up to Reynolds numbers of 240,000 (15 m/s) is explored using a single dielectric barrier discharge (DBD) plasma actuator near the flap shoulder. Results show that the plasma discharge can increase or reduce the size of the time-averaged separated region over the flap depending on the frequency of actuation. High-frequency actuation, referred to here as quasi-steady forcing, slightly delays separation while lengthening and flattening the separated region without drastically increasing the measured lift. The actuator is found to be most effective for increasing lift when operated in an unsteady fashion at the natural oscillation frequency of the trailing edge flow field. Results indicate that the primary control mechanism in this configuration is an enhancement of the natural vortex shedding that promotes further momentum transfer between the freestream and separated region. Based on these results, different modulation waveforms for creating unsteady DBD plasma-induced flows are investigated in an effort to improve control authority. Subsequent measurements show that modulation using duty cycles of 50–70% generates stronger velocity perturbations than sinusoidal modulation in quiescent conditions at the expense of an increased power requirement. Investigation of these modulation waveforms for trailing edge separation control similarly shows that additional increases in lift can be obtained. The dependence of these results on the actuator carrier and modulation frequencies is discussed in detail.

Experimental investigation of a blunt trailing edge flow field with application to sound generation

The unsteady lift generated by turbulence at the trailing edge of an airfoil is a source of radiated sound. The objective of the present research was to measure the velocity field in the near wake region of an asymmetric beveled trailing edge in order to determine the flow mechanisms responsible for the generation of trailing edge noise. Two component velocity measurements were acquired using particle image velocimetry. The chord Reynolds number was 1.9 × 106. The data show velocity field realizations that were typical of a wake flow containing an asymmetric periodic vortex shedding. A phase average decomposition of the velocity field with respect to this shedding process was utilized to separate the large scale turbulent motions that occurred at the vortex shedding frequency (i.e., those responsible for the production of tonal noise) from the smaller scale turbulent motions, which were interpreted to be responsible for the production of broadband sound. The small scale turbulence was found to be dependent on the phase of the vortex shedding process implying a dependence of the broadband sound generated by the trailing edge on the phase of the vortex shedding process.

Radiated sound and turbulent motions in a blunt trailing edge flow field

The dipole sound produced by edge scattering of pressure fluctuations at a trailing edge is most often an undesirable effect in turbomachinery and control surface flows. The ability to model the flow mechanisms associated with the production of trailing edge acoustics is important for the quiet design of such devices. The objective of the present research was to experimentally measure flow field and acoustic variables in order to develop an understanding of the mechanisms that generate trailing edge noise. The results of these experiments have provided insight into the causal relationships between the turbulent flow field, unsteady surface pressure, and radiated far field acoustics. Experimental methods used in this paper include particle image velocimetry (PIV), unsteady surface pressures, and far field acoustic pressures. The model investigated had an asymmetric 45° beveled trailing edge. Reynolds numbers based on chord ranged from 1.2 × 10 6 to 1.9 × 10 6. It was found that the small-scale turbulent motions in the vicinity of the trailing edge were modulated by a large scale von Karman wake instability. The broadband sound produced by these motions was also found to be dependant on the “phase” of the wake instability.

Investigation of flow at leading and trailing edges of pitching-up airfoil

The dynamic stall process of an NACA 0012 airfoil undergoing a constant-rate pitching-up motion is studied experimentally in a water towing tank facility. This study focuses on the detailed measurement of the unsteady separated flow in the vicinity of the leading and trailing edges of the airfoil. The measurements are carried out using the particle image velocimetry technique. This technique provides the two-dimensional velocity and associated vorticity fields, at various instants in time, in the midspan of the airfoil. Near the leading edge, large vortical structures emerge as a consequence of van Dommelen and Shen type separation and a local vorticity accumulation. The interaction of these vortices with the reversing boundary-layer vorticity initiates a secondary flow separation and the formation of a secondary vortex. The mutual induction of this counter-rotating vortex pair eventually leads to the ejection process of the dynamic stall vortex from the leading-edge region. It is found that the trailing-edge flowfield only plays a secondary role on the dynamic stall process.

Experimental investigation of loading effects on compressor trailing-edge flowfields

This article describes an experimental investigation of the flowfield development in the trailing-edge region of a simulated compressor airfoil having pressure-to-suction surface loading. The study represents a continuation of an effort to study turbulent, separated airfoil trailing-edge flows. Previous experimentation had addressed airfoil trailing-edge separation phenomena in the absence of pressure loading. The objective of the current study was to explore the effects of loading and thereby to provide a more realistic simulation of the compressor airfoil flow environment. The present approach was to conduct a large-scale cascade simulation of the airfoil trailingedge flowfield for both a nominal design condition and a higher loading off-design condition. For the design condition, the airfoil boundary layers separated on the blunt trailing edge. For the off-design condition, separation occurred on the suction surface upstream of the trailing edge. A principal result of the study was that pressure loading was found to alter the trailing-edge time mean velocity field from that observed in the previous unloaded airfoil experiment. The relative base pressure for the nominal design condition was unaltered from the unloaded experiment but significantly lower for the off-design condition. Increased loading was found to induce greater cascade exit flow deviation from the exit metal angle, thus, reducing the relative cascade flow turning.

Experimental investigation of a simulated compressor airfoil trailing-edge flowfield

This study was motivated by the need for improved understanding of blunt trailing-edge compressor airfoil flow fields. The objective was to provide data that could be used to assist the development of computational procedures for predicting such flows. A large-scale, low subsonic Mach number, wind tunnel simulation of a compressor airfoil trailing-edge flow field was conducted using a flat-plate test model and laser Doppler velocimetry for flow field definition. Of potential importance in the numerical modeling of the flow was the finding that outer flow velocity profiles in the near wake were nearly identical, with viscous interaction effects confined to a region near the wake centerline having a thickness and length comparable to that of the trailing-edge thickness. Qualitative agreement with circular cylinder near-wake data and an observed lack of dependence of results on a change in plate boundary-layer thickness suggest that the flow is not strongly dependent on boundary-laye r thickness at separation. Periodic vortex shedding was encountered and shown to affect base pressure significantly. The associated strong transverse velocity fluctuations would be expected to be the dominant mechanism for mixing out the velocity defect in the initial wake region.