A Critical Evaluation of Turbulence Modeling in a Model Combustor

Based on the previous benchmark studies on combustion, scalar transfer and radiation models, a critical evaluation of turbulence models in a propane-air diffusion flame combustor with interior and exterior conjugate heat transfers has been performed. Results obtained from six turbulence models are presented and compared in detail with a comprehensive database obtained from a series of experimental measurements. It is found that the Reynolds stress model (RSM), a second moment closure, is superior over the five popular eddy-viscosity two-equation models. Although the main flow patterns are captured by all six turbulence models, only the RSM is able to successfully predict the lengths of both recirculation zones and give fairly accurate predictions for mean velocity, temperature, CO2 and CO mole fractions, as well as turbulence kinetic energy in the combustor chamber. In addition, the realizable k-ε (Rk-ε) model illustrates better performance than four other two-equation models and can provide comparable results to those from the RSM for the configuration and operating conditions considered in the present study.

Flow Control Over a Circular Cylinder Using Pulsed DBD Actuators

Immersed bodies such as struts, vanes, and instrumentation probes in gas turbine flow systems will, except at the lowest of flow velocities, shed separated wakes. These wakes can have both upstream and downstream effects on the surrounding flow. In most applications, surrounding components are designed to be in the presence of a quasi-steady or at least non-variant flow field. The presence of unsteady wakes has both aerodynamic and structural consequences. Active flow control of wake generation can therefore be very valuable. One means to implement active flow control is by the use of plasma actuation. Plasma actuation is the use of strong electric fields to generate ionized gas that can be actuated and controlled using the electric fields. The controlling device can be based on AC, DC, or pulsed-DC actuation. The present research was conducted using pulsed-DC from a capacitive discharge power supply. The study demonstrates the applicability of, specifically, pulsed-DC plasma flow control of the flow on a circular cylinder at high Reynolds numbers. The circular cylinder was selected because its flow characteristics are related to gas turbine flowpath phenomena, and are well characterized. Further, the associated pressure gradients are some of the most severe encountered in fluid applications. The development of effective plasma actuators at high Reynolds numbers under the influence of severe pressure gradients is a necessary step toward developing useful actuators for gas turbine applications beyond laboratory use. The reported experiments were run at Reynolds numbers varying from 50,000 to 97,000, and utilizing various pulse frequencies. Further, the observed performance differences with varying electric field strengths are discussed for these Reynolds numbers. The results show that flow behaviors at high Reynolds numbers can be influenced by these types of actuators. The actuators were able to demonstrate a reduction in both wake width and momentum deficit.

Jet-Surface Interaction Test: Phased Array Noise Source Localization Results

An experiment was conducted to investigate the effect that a planar surface located near a jet flow has on the noise radiated to the far-field. Two different configurations were tested: 1) a shielding configuration in which the surface was located between the jet and the far-field microphones, and 2) a reflecting configuration in which the surface was mounted on the opposite side of the jet, and thus the jet noise was free to reflect off the surface toward the microphones. Both conventional far-field microphone and phased array noise source localization measurements were obtained. This paper discusses phased array results, while a companion paper discusses far-field results. The phased array data show that the axial distribution of noise sources in a jet can vary greatly depending on the jet operating condition and suggests that it would first be necessary to know or be able to predict this distribution in order to be able to predict the amount of noise reduction to expect from a given shielding configuration. The data obtained on both subsonic and supersonic jets show that the noise sources associated with a given frequency of noise tend to move downstream, and therefore, would become more difficult to shield, as jet Mach number increases. The noise source localization data obtained on cold, shock-containing jets suggests that the constructive interference of sound waves that produces noise at a given frequency within a broadband shock noise hump comes primarily from a small number of shocks, rather than from all the shocks at the same time. The reflecting configuration data illustrates that the law of reflection must be satisfied in order for jet noise to reflect off of a surface to an observer, and depending on the relative locations of the jet, the surface, and the observer, only some of the jet noise sources may satisfy this requirement.

Cycle Design Exploration Using Multi-Design Point Approach

The focus of this study is to compare the aerothermodynamic cycle design space of a gas turbine engine generated using two on-design approaches. The traditional approach uses a single design point (SDP) for on-design cycle analysis, where off-design cycle analysis must be performed at other operating conditions of interest. A multi-design point (MDP) method performs on-design cycle analysis at all operating conditions where performance requirements are specified. Effects on the topography of the cycle design space as well as the feasibility of the space are examined. The impacts that performance requirements and cycle assumptions have on the bounds and topography of the feasible space are investigated. The deficiencies of a SDP method in determining an optimum gas turbine engine will be shown for a given set of requirements. Analysis will demonstrate that the MDP method, unlike the SDP method, always obtains a properly sized engine for a set of given requirements and cycle design variables, resulting in an increased feasible region of the aerothermodynamic cycle design space from which the optimum performance engine can be obtained.

An Integrated Approach for Aircraft Engine Performance Estimation and Fault Diagnostics

A Kalman filter-based approach for integrated on-line aircraft engine performance estimation and gas path fault diagnostics is presented. This technique is specifically designed for underdetermined estimation problems where there are more unknown system parameters representing deterioration and faults than available sensor measurements. A previously developed methodology is applied to optimally design a Kalman filter to estimate a vector of tuning parameters, appropriately sized to enable estimation. The estimated tuning parameters can then be transformed into a larger vector of health parameters representing system performance deterioration and fault effects. The results of this study show that basing fault isolation decisions solely on the estimated health parameter vector does not provide ideal results. Furthermore, expanding the number of the health parameters to address additional gas path faults causes a decrease in the estimation accuracy of those health parameters representative of turbomachinery performance deterioration. However, improved fault isolation performance is demonstrated through direct analysis of the estimated tuning parameters produced by the Kalman filter. This was found to provide equivalent or superior accuracy compared to the conventional fault isolation approach based on the analysis of sensed engine outputs, while simplifying online implementation requirements. Results from the application of these techniques to an aircraft engine simulation are presented and discussed.

An Improved Nickel Based MIMS Thermocouple for High Temperature Gas Turbine Applications

A new Nickel based thermocouple for high temperature applications in gas turbines has been devised at the Department of Material Science and Metallurgy of the University of Cambridge. This paper describes the new features of the thermocouple, the drift tests on the first prototype and compares the behaviour of the new sensor with conventional mineral insulated metal sheathed Type K thermocouples: the new thermocouple has a significant improvement in terms of drift and temperature capabilities. Metallurgical analysis has been undertaken on selected sections of the thermocouples exposed at high temperatures which rationalises the reduced drift of the new sensor. A second prototype will be tested in follow-on research, from which further improvements in drift and temperature capabilities are expected.

Total Pressure Correction of a Sub-Miniature Five-Hole Probe in Areas of Pressure Gradients

Five-hole probes, being a dependable and accurate aerodynamic tools, are excellent choices for measuring complex flow fields. However, total pressure gradients can induce measurement errors. The combined effect of the different flow conditions on the ports causes the measured total pressure to be prone to a greater error. This paper proposes a way to correct the total pressure measurement. The correction is based on the difference between the measured total pressure data of a Kiel probe and a sub-miniature prism-type five-hole probe. By comparing them in a ducted fan related flow field, a line of best fit was constructed. The line of best fit is dependent on the slope of the line in a total pressure versus span and difference in total pressure between the probes at the same location. A computer program, performs the comparison and creates the correction equation. The equation is subsequently applied to the five-hole probe total pressure measurement, and the other dependent values are adjusted. The validity of the correction is then tested by placing the Kiel probe and the five-hole probe in ducted fans with a variety of different tip clearances.

Experimental Investigation of the Acoustic Reflection Coefficient of a Modeled Gas Turbine Impingement Cooling Section

The thermoacoustic stability of gas turbines depends on a balance of acoustic energy inside the engine. While the flames produce acoustic energy, other areas like the impingement cooling system contribute to damping. In this paper, we investigate the damping potential of an annular impingement sleeve geometry embedded into a realistic environment. A cold flow test rig was designed to represent real engine conditions in terms of geometry, and flow situation. High quality data was delivered by six piezoelectric dynamic pressure sensors. Experiments were carried out for different mean flow velocities through the cooling holes. The acoustic reflection coefficient of the impingement sleeve was evaluated at a downstream reference location. Further parameters investigated were the number of cooling holes, and the geometry of the chamber surrounding the impingement sleeve. Experimental results show that the determining parameter for the reflection coefficient is the mean flow velocity through the impingement holes. An increase of the mean flow velocity leads to significantly increased damping, and to low values of the reflection coefficient.

Component Map Tuning Procedure Using Adaptive Modeling

Performance models are effective tools for analysis of engine condition throughout the life cycle of a gas turbine engine. Component maps necessary for accurate performance modeling are typically not provided by the original equipment manufacturers. To compensate for the missing information, available maps of similar components are scaled to match component performance at one or more reference points. Although scaled maps can provide sufficiently accurate results close to the reference points, modeling errors tend to increase further away from these reference points. For applications such as gas path analysis, the resulting modeling errors can be of the same order of magnitude as the deterioration to be detected. This severely limits the application of such techniques. This article presents a component map tuning procedure that tunes maps with more detail than just scaling. The tuned maps are a closer match to real component performance. The tuning procedure combines the adaptive modeling capability of the Gas turbine Simulation Program (GSP) and on-wing measured engine performance data. On-wing measured engine performance data allows map tuning over a wider range of power settings compared to engine performance data measured in a test cell. Effects of measurement uncertainty and scatter, and effects of compressor bleed flows on the map tuning procedure are analyzed and discussed. The tuned component maps enabled more accurate component condition estimations, mainly characterized by less scatter. By improving the accuracy of gas path analysis with on-wing measured performance data, this work has enabled more effective use of performance diagnostic techniques in the aero-engine maintenance industry.

Electro-Thermal Measurement of Heat Transfer Coefficients

Heat transfer coefficients are very important for the design of the various flow paths found in turbomachinery. Therefore, the measurement of heat transfer coefficients plays an important role in the field of turbomachinery research. An accurate measurement of heat transfer is not a simple task considering gaseous flow in combination with good thermal conductivity of the boundaries along the flow path. The majority of the measurement methods applied has at least one of the following problems. The measurement setup as for instance a heat flux sensor is a thermal barrier in the object of interest or the sensor introduces for measurement reasons a lot of heat into the object of interest. In both cases the main error results from the modification of the system, which is critical for the investigation of any kind of flow influenced by buoyancy. Furthermore, insufficient fluid reference temperature and/or heat flux with changing sign corrupts any attempt to calculate reasonably heat transfer coefficients. The measurement of heat transfer coefficients becomes even more complicated if the flow path of interest rotates at some thousand rpm as for instance in gas turbines or any other fast rotating machine with fluid flow. This contribution presents an experimental investigation of a setup for the direct measurement of heat transfer coefficients in gaseous flow with metallic boundaries. The calibration of a test sample probe is presented for the standard case forced convection on a flat plate. The sensor setup provides low influence of the measurement on the object investigated. It overcomes the problem of a reference temperature and delivers always positive heat transfer coefficients. Furthermore, the measurement setup fulfils the requirements of telemetric application in the rotating system of a machine. The test of the sensor for its strength against centrifugal acceleration is part of the continuation of the work.