Ephraim Gutmark

Large-Eddy Simulations of a Supersonic Jet and Its Near-Field Acoustic Properties

Large-eddy simulations of imperfectly expanded jet flows from a convergent-divergent nozzle with a sharp contraction at the nozzle throat have been carried out. The flowfield and near-field acoustics for various total pressure ratios from overexpanded to underexpanded jet flow conditions have been investigated. The location and spacing of the shock cells are in good agreement with experimental data and previous theoretical results. The velocity profiles are also in good agreement with data from experimental measurements. A Mach disk is observed immediately downstream of the nozzle exit for overexpanded jet conditions with nozzle pressure ratios much lower than the fully expanded value. It is found that this type of nozzle with a sharp turning throat does not have a shock-free condition for supersonic jet flows. The near-field intensities of pressure fluctuations show wavy structures for cases in which screech tones are observed. The large-eddy simulations predictions of the near-field noise intensities show good agreement with those obtained from experimental measurements. This good agreement shows that large-eddy simulations and measurements can play complementary roles in the investigation of the noise generation from supersonic jet flows.

Supersonic Jet Noise Reduction Technologies for Gas Turbine Engines

This paper presents observations and simulations of the impact of several technologies on modifying the flow-field and acoustic emissions from supersonic jets from nozzles typical of those used on military aircraft. The flow-field is measured experimentally by shadowgraph and particle image velocimetry. The acoustics are characterized by near- and far-field microphone measurements. The flow- and near-field pressures are simulated by a monotonically integrated large eddy simulation. Use of unstructured grids allows accurate modeling of the nozzle geometry. The emphasis of the work is on “off-design” or nonideally expanded flow conditions. The technologies applied to these nozzles include chevrons, fluidic injection, and fluidically enhanced chevrons. The fluidic injection geometry and the fluidic enhancement geometry follow the approach found successful for subsonic jets by employing jets pitched 60 deg into the flow, impinging on the shear layer just past the tips of the chevrons or in the same axial position when injection is without chevrons.

Experimental Study of a Multinozzle Combustor at Elevated Pressures

Emissions measurements and OH* chemiluminescence imaging were performed on a novel multipoint lean direct injection combustor designed for low-NOx emissions and with the flexibility to operate at both high- and low-power settings. The combustor has five rows of fuel nozzles that can be staged in three independent circuits as required. The fuel nozzle stages are not identical; instead, they are optimized for different power settings. NOx, CO, and unburned hydrocarbons were measured for a variety of settings, including lean blowout, idle, intermediate power, and simulated full power. Effects of pressure, inlet temperature, equivalence ratio, and fuel flow distribution were studied. OH* chemiluminescence was used to observe flame structure and interpret emissions trends. The local equivalence ratio, calculated with the fuel and air flow rates particular to each combustor stage, was a useful parameter for interpreting emissions. The combustor demonstrated the ability to operate at a wide range of conditions, ...

Characterization of Shock Wave Transfer in a Pulse Detonation Engine–Crossover System

Shock wave propagation within a pulse detonation engine–crossover system is investigated, examining the properties and mechanisms of the transfer process. A shock wave is transferred through a crossover tube that connects a spark-ignited driver pulse detonation engine to a secondary, driven pulse detonation engine. Detonations in the driven pulse detonation engine develop from shock-initiated combustion, as strong shock wave reflection can cause ignition within a reactive mixture. A pulse detonation engine–crossover system can decrease deflagration-to-detonation transition length while employing a single spark source to initiate a system consisting of multiple detonation engines. Visualization of a shock wave propagating through a clear channel reveals a complex shock train behind the leading shock wave. Transverse waves connect with the leading shock wave to form a triple point that oscillates through the leading shock wave. The shock wave Mach number and decay rate remain constant for varying crossover ...

Experimental Comparison of Axial Turbine Performance Under Steady and Pulsating Flows

The performance of an axial turbine is studied under close-coupled, out-of-phase, multiple-admission pulsed air flow to approximate turbine behavior under pulsed detonation inflow. The operating range has been mapped for four frequencies and compared using multiple averaging approaches and five formulations of efficiency. Steady performance data for full and partial admission are presented as a basis for comparison to the pulsed flow cases. While time-averaged methods are found to be unsuitable, mass-averaged, work-averaged, and integrated instantaneous methods yield physically meaningful values and comparable trends for all frequencies. Peak work-averaged efficiency for pulsed flow cases is within 5% of the peak steady, full admission values for all frequencies, in contrast to the roughly 15–20% performance deficit experienced under steady, 50% partial admission conditions. Turbine efficiency is found to be a strong function of corrected flow rate and mass-averaged rotor incidence angle, but only weakly dependent on frequency.Copyright

Supersonic turbojet noise reduction

Observations and Large Eddy Simulations are presented of a supersonic jet from a nozzle representative of high-performance military aircraft such as the Saab Gripen. The nozzle has a design Mach number of 1.56 and is examined at its design condition with a surrounding secondary flow at Mach numbers of 0.0, 0.1 and 0.3. The nozzle is investigated in its unmodified state and also with the addition of chevrons and microjets. Detailed flow-field velocity measurements of the jets and far-field noise measurements are presented and the noise results are scaled to represent the effects of the chevrons and microjets on airport neighbors. Chevrons and internal fluidic injection by microjets each reduce the noise generated by the main jet. And substantially reduce the noise footprint around the airport. The numerical simulation technique, correctly predicts the flow and noise not only the baseline case, but also the noise reduction by both chevrons and microjects.

Experimental and Numerical Study of Jets from Elliptic Nozzles with Conic Plug

The study presents results concerning jets exhausting from elliptic nozzles with conic centerbody plugs. A nozzle with 3:1 ratio of exit heights is surveyed experimentally using particle image velocimetry for jet Mach numbers from 0.24 to 1.0, heated and unheated. At the exit, the nozzle inner surface has a slope of zero in the major axis plane, while in the minor axis plane the slope is steeper than that for a corresponding round nozzle. The conic plug causes splitting of the initial elliptic jet into two jets with their centerlines in the major-axis plane of the elliptic nozzle. Increasing M j or total temperature causes the bifurcated potential cores of the individual jets to be slightly smaller and to diverge from one to another at a slightly greater angle from the nozzle centerline. Large eddy simulations of the experimental nozzle and two alternate nozzles are analyzed to find how the flow features are influenced by not forcing the jet towards the conic plug in the minor axis plane and by changing the ratio of exit heights to 2:1. Changing nozzle shape caused reduction in the spreading rate in the major axis and increase in spreading in the minor axis. Bifurcation occurs for all nozzles studied.