A team led by General Electric Research (GER) and Southwest Research Institute (SwRI) was tasked to design, build and test an advanced 4MW CO2 compressor that would operate near the liquid-vapor dome for Carbon Dioxide (CO2). The US Department of Energy (DoE) Solar Technologies Office (SETO) funded program was targeted towards a Concentrated Solar Power (CSP) plant where optimum power cycle efficiency can be obtained when operated close to the liquid-vapor dome where CO2 is a supercritical fluid (sCO2) as compression power is reduced in the main compressor. However, the CSP cycle and other related supercritical CO2 cycles (fossil, nuclear, waste heat recovery) have considerable compression challenges both mechanically and aerodynamically when operating with a high density fluid that exceeds 70% the density of water. The subject of this paper is highlighting the challenge in determining compressor performance using industry standard measurements. This application is the highest density industrial-scale centrifugal compressor in the world at 720 kg/m 3 . This paper will investigate the uncertainty when measuring compressor efficiency using ASME PTC-10 instrumentation and the effect of the strong CO2 property variation when operating as a supercritical fluid, near the fluid-vapor dome. Prior work in this area by Wahl will be summarized and compared with the current compressor test program uncertainty. It will be shown that Wahl predicted high uncertainty as well although, the current testing program is even closer to the liquid-vapor dome than the test program under Wahl. The uncertainty analysis has shown that traditional PTC-10 temperature measurements lead to high levels of uncertainty for sCO2 compression near the liquid-vapor dome. The uncertainty is driven by the large changes in thermodynamic properties of sCO2. These property changes are affected by the measured pressure and temperature; however, temperature measurement error is the primary contributor to uncertainty. Because of this, looking at alternate sCO2 property measurements was investigated. Higher quality localized pressure calibration, improving flow measurement accuracy, and measuring density in addition to temperature all significantly improved efficiency uncertainty. The authors confirmed the most significant measurement change is to measure pressure and density through either a densitometer or a Coriolis flow meter which provides a density measurement in conjunction with flow rate accuracy.

The aerodynamic effects of a probe for stage performance evaluation in a high-speed axial compressor are investigated. Regarding the probe measurement accuracy and its aerodynamic effects, the upstream/downstream effects on the probe and probe insertion effects are studied by using an unsteady computational fluid dynamics (CFD) analysis and by verifying in two types of multistage high-speed axial compressor measurements. The probe traverse measurements were conducted at the stator inlet and outlet in each case to evaluate blade row performance quantitatively and its flow field. In the past study, the simple approximation method was carried out which considered only the interference of the probe effect based on the reduction of the mass flow by the probe blockage for the compressor performance, but it did not agree well with the measured results. In order to correctly and quantitatively grasp the mechanism of the flow field when the probe is inserted, the unsteady calculation including the probe geometry was carried out in the present study. Unsteady calculation was performed with a probe inserted completely between the rotor and stator of a 4-stage axial compressor. Since the probe blockage and potential flow field, which mean the pressure change region induced by the probe, change the operating point of the upstream rotor and increase the work of the rotor. Compared the measurement result with probe to a kiel probe setting in the stator leading edge, the total pressure was increased about 2,000Pa at the probe tip. In addition, the developed wake by the probe interferes with the downstream stator row and locally changes the static pressure at the stator exit. To evaluate the probe insertion effect, unsteady calculations with probe at three different immersion heights at the stator downstream in an 8-stage axial compressor are performed. The static pressure value of the probe tip was increased about 3,000Pa in the hub region compared to tip region, this increase corresponds to the measurement trend. On the other hand, the measured wall static pressure showed that there is no drastic change in the radial direction. In addition, when the probe is inserted from the tip to hub region in the measurement, the blockage induced by the probe was increased. As a result, operating point of the stator was locally changed, and the rise of static pressure of the stator increased when the stator incidence changed. These typical results show that unsteady simulations including probe geometry can accurately evaluate the aerodynamic effects of probes in the high-speed axial compressor. Therefore, since the probe will pinpointed and strong affects the practically local flow field in all rotor upstream passage and stator downstream, as for the probe measurement, it is important to pay attention to design the probe diameter, the distance from the blade row, and its relative position to the downstream stator. From the above investigations, a newly simple approximation method which includes the effect of the pressure change evaluation by the probe is proposed, and it is verified in the 4-stage compressor case as an example. In this method, the effects of the distance between the rotor trailing edge (T.E.) and the probe are considered by the theory of the incompressible two-dimensional potential flow. The probe blockage decreases the mass flow rate and changes the operating point of the compressor. The verification results conducted in real compressor indicate that the correct blockage approximation enables designer to estimate aerodynamic effects of the probe correctly.

This publication deals with the improvement of the design and life steam testing of a low-pressure steam turbine module for utilization in industrial applications. Extensive numerical investigations were carried out and validated with experimental measurements. For the validation of the final blade design, the LP module was integrated and tested in a 12.5MW steam turbine test rig. During the balancing of the rotor in a vacuum chamber, tip-timing and strain gage measurements were conducted for testing the mechanical behavior of the airfoils. In the succeeding assembly of the steam turbine, tip-timing probes and flow measurement probes were integrated around the LP module to evaluate the blade load and performance under realistic conditions. To obtain a good understanding of the improved performance and mechanical behavior the thermodynamic tests were carried out under no-load, part load and design load conditions to cover the whole operating range of the LP stages. The paper also describes how an existing rotor from the test rig could be retrofitted, to accommodate the new turbine stages with additive manufacturing of the base material on the shaft.

Given the robustness and maturity of contemporary particle image velocimetry (PIV) methods, detailed flow measurements are now possible in actual turbine engine environments. For instance, flow non-uniformity measurements are possible in fan outlet gas-paths by adapting PIV hardware to the restricted access afforded in such applications. In the present work, a framework is proposed and demonstrated for planning and executing stereoscopic PIV measurements in the fan outlet duct of a Pratt & Whitney Canada JT15D-1A research turbofan engine. Two case studies have been carried out by following this framework in order to demonstrate two different imaging methods — conventional lens/camera coupling and endoscopic imaging. A key step within the planning and execution framework is risk reduction, and this step resulted in considerable refinement of the methods in both cases. For instance, in endoscopic imaging, the risk reduction provided a new fluorescent paint application for obtaining near wall data not previously possible. The data obtained in the engine experiments have been used to quantify the uncertainty for both imaging methods, not surprisingly revealing 50% greater average uncertainties for endoscopic PIV versus conventional lens/camera imaging. The results from the case studies indicate that both imaging methods may be practically and economically implemented for detailed measurements in fan outlet ducts, and application demands may be carefully considered, and risks reduced, using the framework proposed.

A three-dimensional flow in a high speed (Ma in = 0.7) linear compressor cascade’s end wall region is presented in this paper. The attached blade geometry is of medium curvature with a deflection angle Δβ = 15°. To intensify the secondary flow structure the profiles were stalled deliberately and oil flow visualization was used to prove the presence of that hub-corner stall. Together with streamlines derived from velocities, measured by a three-channel laser Doppler anemometer (LDA), that allowed the identification of the secondary flow region’s key features and their extents into the flow field. A therefrom-derived topology is then compared to existing concepts. Moreover, an associated URANS simulation is set up using the k–ε model to verify its capabilities to reproduce the measured flow field. To confirm the previous observations the stall indicator is computed from the simulation results. It is further shown that the hub-corner stall dissolves when removing the side wall boundary layer. In summary, this paper combines elaborate laser Doppler measurements at both high Reynolds and Mach numbers combined with verified simulations allowing a detailed description of the end wall behavior within the cascade and further introduces a new flow topology concept.

The rapid regulation of gas turbine power plants is becoming increasingly important. In the field of stationary gas turbine power augmentation droplet-laden flows receive special attention as they cool the compressor’s inlet temperature. High fogging is an effective and economic method that secures power supply by reacting rapidly to fluctuating power generation. However, water droplets effect the aerodynamic performance of the compressor negatively. Therefore, the Laboratory of Turbomachinery investigated the interaction between droplets and blades at midspan by using diverse experimental methods to comprehend aerodynamic performance and water film induced losses. Additionally, the knowledge of the separation’s extent for three-dimensional flows with corner separation is of great importance as it leads to a blockage of the flow cross-section and therefore contributes to the losses. This paper enlarges this knowledge base by studying the interactions of the dispersed phase with the gas flow in three-dimensional flow structures as they are found in the near wall regions of every turbo compressor. By utilizing a three-dimensional Phase-Doppler Anemometer spatially resolved information on velocity mean vector, its higher moments as well as droplet diameters were extracted from the separated region for a compressor blade, in middle cross section design, with spray-injection. Streamlines that derive from the mean velocities visualize key features such as saddle and focal points. Then an integral approach is presented to evaluate the effect of the dispersed phase on the flow structure compared to dry flow and the impact back on different diameter classes. Finally, the velocity in streamwise direction shows the influence of the disperse phase on the aerodynamic blockage.

Performance and flow measurements are carried out to investigate the impact of varying the geometry of axial casing grooves on the stall margin and efficiency of an axial turbomachine. Prior studies have shown that skewed semi-circular grooves installed near the blade leading edge (LE) have multiple effects on the flow structure, including ingestion of the tip leakage vortex (TLV), suppression of backflow vortices, and periodic variations of flow angle. To determine which of these phenomena is a key contributor, the present study examines the impact of several grooves, all with the same inlet geometry, but with outlets aimed at different directions. The “U” grooves that have circumferential exits aimed against the direction of blade rotation achieve the highest stall margin improvement of well above 60% but cause a 2.0% efficiency loss near the best efficiency point (BEP). The “S” grooves, which have exits aimed with the blade rotation, achieve a relatively moderate stall margin improvement of 36%, but they do not reduce the BEP efficiency. Other grooves, which are aligned with and against the flow direction at the exit from upstream inlet guide vanes, achieve lower improvements. These trends suggest that causing high periodic variations in flow angle around the blade leading edge is particularly effective in extending the stall margin, but also reduces the peak efficiency. In contrast, maintaining low flow angles near the LE achieves more moderate improvement in stall margin, without the maximum efficiency loss. Hence, of the geometries tested, the S grooves appear to have the best overall impact on the machine performance. Velocity measurements and flow visualizations are performed in an axial plane located downstream of the grooves, near the trailing edge of the rotor. Reduced efficiency or performance co-occurs with elevated circumferential velocity in the tip region, but differences in the axial blockage are subtle. Yet, near the BEP, the regions with reduced axial velocity, or even negative velocity between the TLV and the endwall, are wider behind the U grooves compared to the S grooves. The vorticity profiles also show that at low flow rates the TLV is ingested entirely by the grooves, in contrast to the best efficiency point, where a considerable fraction of the TLV rollup occurs downstream of the grooves.

Understanding the amount of air that enters a gas turbine is important for calculating their performance and efficiency. Flow meters are almost never used to measure that flow in production engines. Typical flow meters are impractical because the air flow into the compressor is very large, up to 1400 lbs/s (635 kg/s) or 1,000,000 ft 3 /min (28,300 m 3 /min), and typically an intentional pressure drop is involved in the measurement. This pressure drop negatively impacts the performance of the engine. If inlet air mass flow were known accurately without negatively impacting the engine performance, then engines could be run more efficiently. Currently, inlet mass flow is typically inferred, rather than measured. This leads to increased safety margins which require engines to be run more conservatively, i.e., at lower power. This paper describes a novel, inexpensive, and accurate air mass flow measurement system with negligible impact on engine performance.

This paper presents developing secondary flow and heat transfer measurements in a ribbed cooling channel. Experiments are carried out for Reynolds number ranging from 25,000–140,000. Regionally averaged local heat transfer measurements are conducted using heated copper segments. Flow measurements are carried out using a miniature five-hole pressure probe and presented for cross sections at intervals of 1.8 hydraulic diameters d h in flow direction. Results are compared to numerical simulations using explicit algebraic Reynolds stress and turbulent heat transfer models. The paper focuses on the entrance region where secondary flow structure has not emerged yet. The findings show that the well-known secondary flow structure of the crossed rib configuration, consisting of one large single rotating secondary flow, is not established until approximately 6–7 d h in main flow direction. Instead two opposed vortices are identified which dominate the flow characteristics and provide an increase in heat transfer of up to 15–20% when compared to the periodically developed flow condition. Thus, for the first time to the author’s knowledge, the paper describes in detail the developing secondary flow in a crossed rib arrangement and links it to the heat transfer distribution observed. In summary, this paper stresses the importance of the developing flow region for the design process in convection cooled gas turbines, especially for short channels of high pressure blades and vanes, as it has a significant effect on cooling channel heat transfer performance.

This paper describes testing an axial turbine response when driven by a Rotating Detonation Combustor (RDC). A T63 (C20-250) gas turbine is modified by replacing the combustor with a RDC. The stator vanes of the T63 are heavily instrumented for measurement of flow enthalpy and pressure. The engine is run at multiple power levels with the stock combustor using JetA and hydrogen fuel. The engine is then modified to have an open loop configuration, and is run with both the RDC and the stock combustor hardware with hydrogen fuel. Temperature pattern factor, flow unsteadiness, and turbine component efficiency are measured for all setups. High speed pressure transducers show substantially higher unsteadiness generated by the RDC than the conventional combustor. RDC turbine component efficiencies are compared to the conventional combustor. Results suggest that RDC unsteadiness does not significantly impact turbine efficiency.

Impingement cooling in airfoils cooling cavities, solely or combined with film and convective cooling, is a common practice in gas turbines. Depending on the cooling cavity design, the mass flow rate through individual crossover holes could vary significantly in the flow direction thus creating jets of different strengths in the target cavity. This jet flow variation, in turn, creates an impingement heat transfer coefficient variation along the channel. A test section, simulating two adjacent cooling cavities on the trailing side of an airfoil, is made up of two channels with trapezoidal cross-sectional areas. On the partition wall between the two channels, eleven crossover holes create the jets. Two distinct exit flow arrangements are investigated — a) jets, after interaction with the target surface, are turned towards the target channel exit axially and b) jets are exited from a row of racetrack-shaped slots along the target channel. Flow measurements are reported for individual holes and heat transfer coefficients on the eleven target walls downstream the jets are measured using the steady-state liquid crystal thermography technique. Smooth as well as rib-roughened target surfaces with four rib geometries (0°,45°, 90° and 135° rib angles) are tested. Correlations are developed for mass flow rate through each crossover hole for cases with different number of crossover holes, based on the pressure drop across the holes. Heat transfer coefficient variations along the target channel for all rib geometries and flow conditions are reported for a range of 5000 to 50000 local jet Reynolds numbers. Major conclusions of this study are: 1) A correlation is developed to successfully predict the mass flow rates through individual crossover holes for geometries with six to eleven crossover holes, based on the pressure drop across the holes, 2) impingement heat transfer coefficient correlates well with the local jet Reynolds number for both exit flow arrangements, and 3) the case of axial flow in the target channel exiting from the channel end, at higher jet Reynolds numbers, produced higher heat transfer coefficients than those in the case of flow exiting through a row of slots along the target channel opposite to the crossover holes.

Accurate temperature prediction of turbine blades for gas turbine is very important to assure the life-span of the blade under a hostile hot gas environment and intense centrifugal force. Therefore, there have been a number of studies carried out to clarify the cooling performance of serpentine cooling channel inside a turbine blade for gas turbine, however, it remains to be quite difficult to make an accurate numerical prediction of the performance. Apart from the effects of disk rotation as well as large temperature gradient near the wall, such a poor predictability can be attributed to the complicated vortical motions caused by the rib-roughened cooling channel whose cross-sectional shape varies along the channel and by the existence of u-bends. Furthermore, since the cooling channel inside a real turbine blade usually has a curved or S-shaped inlet, which may induce flow separation as well as swirl developed in the inlet, it can be imagined that the flow and heat transfer inside the cooling channel is likely to become much more complicated than that with a straight inlet. Despite this situation, only few studies are made in order to examine the flow and heat transfer characteristics inside the cooling channel with s-shaped inlet. Accordingly, this study aims at detailed experimental and numerical investigations on the flow and heat transfer characteristics of a realistic serpentine rib-roughened cooling channel with an s-shaped inlet, which is modeled from an actual HP turbine blade for gas turbine. This study employs a transient TLC (Thermochromic Liquid Crystal) technique to measure the heat transfer characteristics, along with the flow visualization on the inner surface of the channel using oil mixed with titanium powder. Note that a special focus in this flow visualization is placed on the area of s-shaped inlet. As for the flow measurement, 2D-PIV (Particle Image Velocimetry) method is used to understand time-dependent vortical structures of the flow field that can have significant impacts on the heat transfer. RANS-based numerical simulation is also executed to predict the heat transfer distribution on the inner surface of the cooling channel.

Printed circuit heat exchangers (PCHEs) have an important role in supercritical CO 2 (sCO 2 ) Brayton cycles because of their small footprint and the high level of recuperation required for this power cycle. Compact heat exchangers like PCHEs are a rapidly evolving technology, with many companies developing various designs. One technical unknown that is common to all compact heat exchangers is the flow distribution inside the headers that affects channel flow uniformity. For compact heat exchangers, the core frontal area is often large compared with the inlet pipe area, increasing the possibility of flow maldistribution. With the large area difference, there is potential for higher flow near the center and lower flow around the edges of the core. Flow maldistribution increases pressure drop and decreases effectiveness. In some header geometries, flow separation inside the header adds to the pressure drop without increasing heat transfer. This is the first known experiment to test for flow maldistribution by direct velocity measurements in the headers. A PCHE visualization prototype was constructed out of transparent acrylic for optical flow measurements with Particle Image Velocimetry (PIV). The channels were machined out of sheets to form many semi-circular cross sections typical of chemically-etched plates used in PCHE fabrication. These plates were stacked and bolted together to resemble the core geometry. Two header geometries were tested, round and square, both with a normally-oriented jet. PIV allows for velocities to be measured in an entire plane instantly without disturbing the flow. Small particles of approximately 10 micrometers in diameter were added to unheated water. The particles were illuminated by two laser flashes that were carefully timed, and two images were acquired with a specialized digital camera. The movement of particle groups was detected by a cross-correlation algorithm with a result of about 50k velocity measurements in a plane. The velocity distribution inside the header volume was mapped using this method over many planes by traversing the PCHE relative to the optical equipment. The level of flow maldistribution was measured by the spatially-changing velocity coming out of the channels. This effect was quantified by the coefficient of variation proposed by Baek et al. The relative levels of flow maldistribution in the different header geometries in this study were assessed. With highly-resolved velocity measurements, improvements to header geometry to reduce flow maldistribution can be developed.

A robust 1D film hydrodynamic model has been sequentially coupled with a 1D core gas model and used to predict the instantaneous mean core gas speed, film interface shear stress and liquid film distribution within an idealised bearing chamber. This novel approach to aero-engine bearing chamber simulation provides a predictive tool that can be used for the fast and reliable exploration of a set of bearing chamber design and operating conditions characterised by the: chamber dimensions, air/oil fluid properties, shaft speed, sealing air flows, oil feed rates and sump scavenge ratios. A preliminary validation of the model against available bearing chamber flow measurements from literature shows good agreement. The model represents a significant step change in predictive capabilities for aero-engine oil system flows compared to previous semi-empirical models. The bearing chamber is idealised as a one-dimensional (2D) domain with a predominantly azimuthal flow in both the rotational oil film and core gas such that axial components may be ignored. A 1D system of depth-averaged film hydrodynamics equations is used to predict oil film thickness and mean speed distributions in the azimuthal direction under the influence of interface shear, gravity, pressure gradient and surface tension forces. The driving shear stress in the film model is obtained from the 1D core-gas model based on an azimuthal gas momentum conservation equation which is coupled to the film model through the interface shear stress and film interface velocity.

A rotating test rig, reproducing a rotor-stator cavity with an axial admission system, has been exploited for an experimental investigation on the internal flow field and its effect on heat transfer on the stator side. Working conditions were varied in a wide range of rotating velocities and superposed mass flow rates. 2D PIV flow measurements were performed in order to obtain a radial distribution of the tangential velocity, results were used to validate numerical simulations aimed at understanding the admission system effect on the swirl distribution. Heat transfer coefficient distribution along the stator disk has been evaluated performing a steady state technique exploiting Thermo-chromic Liquid Crystals (TLC). Tests have been performed varying the superposed mass flow rate up to reaching the condition of cavity completely sealed, further increase of the mass flow rate showed to reduce the effect of the rotation. Working conditions were set in order to investigate cases missing in open literature, however few tests performed in similarity with other researches provided comparable results.

Currently fast response aerodynamic probes are widely used for advanced experimental investigations in turbomachinery applications. The most common configuration is a virtual three hole probe. This solution is a good compromise between probe dimension and accuracy. Several authors have attempted to extend the capabilities of these probes in terms of bandwidth and operating conditions. Even though differences exist between the solutions in the literature, all of the designs involve the positioning of a dynamic pressure sensor close to the measurement point. In general terms, the higher the frequency response, the more the sensor is exposed to the flow. This physical constraint puts a limit on the probe applicability since the measurement conditions have to comply with the maximum allowed operating conditions of the sensor. In other applications, when the conditions are particularly harsh and a direct measurement is not possible, a waveguide probe is commonly used to estimate the local pressure. In this device the sensor is connected to the measurement point through a transmitting duct which guarantees that the sensor is operating in a less critical condition. Generally, the measurement is performed through a pressure tap and particular attention must be paid to the probe design in order to have an acceptable frequency response function. In this study, the authors conceived, developed and tested a probe which combines the concept of a fast response aerodynamic pressure probe with that of a waveguide probe. Such a device exploits the benefits of having the sensor far from the harsh conditions while maintaining the capability to perform an accurate flow measurement.

This paper presents an analytical approach to quantify the uncertainty of velocity-, angle- and turbulence intensity measurements obtained by hot-wire anemometry. Based on the given results of the model, the paper points out the most critical parameters influencing the quality of the flow measurement. The proposed approach is divided into three subsections, each accounting for the uncertainties due to the measurement equipment, probe calibration and measurements in the test-rig respectively. For every subsection the propagation of uncertainties is stated in the form of analytical formulas starting with the measurands, e.g. voltage, pressure, temperature, and ending with velocity, angle and turbulence intensity. The approach is applied to data, obtained by a triple hot-wire probe, on two-dimensional traverses in a 1.5-stage cold air turbine. The calibration process and, in particular, the temperature and pressure deviation between calibration and measurement are identified as having a major impact on the measurement uncertainty. With respect to the Aachen test-rig the uncertainty in velocity measurements is lower than 3.5%. The investigation of the turbulence intensity uncertainty yields an increase in uncertainty with increasing turbulence intensity but an insensitivity towards pitch- and yaw-angle uncertainty.

The full-field endwall Nusselt number ( Nu ) distributions and flow field are presented respectively using steady-state infrared thermography and particle image velocimetry (PIV) for the two-pass parallelogram channels with attached and detached transverse ribs. These square transverse ribs on two opposite channel endwalls are in-line arranged with rib-height to duct-height ratio of 0.1 and rib-pitch to rib-height ratio of 10. For the detached ribs, the detached distance between rib and channel endwall is 0.38 rib height. With the measurements of Fanning friction factor ( f ), heat transfer distributions and flow field features, the thermal performance factors (TPF) for the attached and detached rib cases are comparatively examined. A set of Nu , f and TPF results with the associated flow measurements at the test conditions of 5,000≤ Re ≤20,000 is selected to disclose the differential heat transfer enhancement mechanisms and heat transfer efficiencies between the attached and detached ribbed channels. Empirical correlations evaluating the endwall area-averaged Nusselt numbers ( Nu ) and f factors are devised to assist the relevant design activities.

Numerical simulation results of a single jet premixed combustion system at atmospheric pressure are compared against comprehensive particle image velocimetry (PIV) flow measurements and Raman scattering temperature measurements for natural gas and hydrogen fuels. The simulations were performed on hexahedral meshes with 1–5 million elements. RANS calculations were carried with the k-ε realizable turbulence model. Combustion was modeled using the Flamelet Generated Manifold model (FGM) and detailed chemistry. Both the flame position and flame liftoff predicted by the FGM were in reasonable agreement with experiments for both fuels and showed little sensitivity to heat transfer or radiation modeling. The detailed chemistry calculation predicts the temperature gradients along the jet centerline accurately and compares very closely with the Raman scattering measurements. The much closer agreement of the jet axial velocity and temperature profiles with experimental values, coupled with the significantly protracted presence of intermediates in the detailed chemistry predictions, indicates that the impact of nonequilibrium intermediates on very lean natural gas flames is significant.

In this study, we experimentally investigate both the intrinsic instability characteristics and forced response to transverse acoustic excitation of a non-reacting, swirling flow for application to combustion instability in annular gas turbine engines. The non-axisymmetry of the velocity field is quantified using an azimuthal mode decomposition of the time-averaged velocity field that shows that (1) the flow field is largely axisymmetric, (2) axisymmetry decreases with downstream distance, and (3) forcing does not significantly alter the time-averaged shape of the flow field. The flow field is analyzed in a companion linear stability analysis that shows that the most unstable modes in the flow field are m =−1 and m =−2, which agrees with the experimental observations and shows that the intrinsic dynamics of this flow field are non-axisymmetric with respect to the jet axis. The linear stability analysis captures the spatial variation of mode strength for certain modes, particularly mode m =−1, but there are some deviations from the experimental results. Most notably, these deviations occur for mode m =0 at radii away from the jet axis. Experimental results of the forced response of the flow indicate that the intrinsic instability characteristics of the flow field have an impact on the forced-response dynamics. Response of the flow field to a velocity anti-node in a standing transverse acoustic field shows non-axisymmetric vortex rollup and the dominance of the m =−1 and m =1 azimuthal modes in the fluctuating flow field. In the presence of a pressure anti-node, the m =0 mode of the fluctuating flow field is very strong at the jet exit, indicating an axisymmetric response, and ring vortex shedding is apparent in the flow measurements from high-speed PIV. However, further downstream, the strength of the axisymmetric mode decreases and the m =−1 and m =1 modes dominate, resulting in a tilting of the vortex ring as it convects downstream. Implications for flame response to transverse acoustic fields are discussed.