David Munday

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.

Flow structure and acoustics of supersonic jets from conical convergent-divergent nozzles

Conical convergent-divergent (CCD) nozzles represent an important category of supersonic jet-engine nozzles which require variable throat areas and variable exit areas to adapt to a range of operating conditions. CCD nozzles with design Mach numbers of 1.3, 1.5, and 1.65 are examined experimentally over a range of fully expanded Mach numbers from 1.22 to 1.71. The characteristics of the flow and acoustic fields from these nozzles are explored. Shadowgraph, Particle Image Velocimetry, far-field and near-field acoustic surveys are presented. Results of a Monotonically Integrated Large Eddy Simulation are presented for the Mach 1.5 nozzle at an underexpanded condition. The agreement between simulations and measurements is excellent. It is shown that these nozzles differ from traditional smoothly contoured method-of-characteristics nozzles in that they never achieve a shock free condition. Furthermore it is shown that these nozzles produce a “double diamond” pattern in which two sets of shock diamonds are generated with an axial displacement between them. The cause of this phenomenon is explored. It is further shown that as a consequence they are never free from shock-associated noise even when operated at perfect expansion. In spite of this difference, it is found that CCD nozzles behave like traditional convergent-divergent nozzles in that they produce the same shock-cell size, broadband shock-associated noise peak frequency, and screech frequency as traditional convergent-divergent nozzles. The apparent source regions for mixing noise, broadband shock associated noise and screech are all similar to those from traditional convergent-divergent nozzles.

ACTIVE CONTROL OF SEPARATION ON A WING WITH CONFORMAL CAMBER

A preliminary investigation of a wing with a conformal camber is discussed. The wing uses an adaptive actuator mounted internally to alter the shape of the suction surface which results in a change in the effective camber by increasing the maximum thickness and moving the location of maximum thickness aft. Since the actuator motion can be altered continuously, this allows the wing shape to be either static or dynamic. For the static mode, effects of profile perturbations are tested in a wind tunnel and tow tank to determine the change in wing efficiency due to variations in the camber at low Reynolds number. Various actuator locations and frequencies are tested from Re = 2.5 · 10 4 to Re = 2 · 10 5 at varying angles of attack. Preliminary comparisons are made with numerical predictions to demonstrate the effect of adaptive wing shaping on flight performance and the difficulty of predicting separated flow at low Re. For the dynamic mode, a wing with 5 modular span-wise sections each with a separately controlled internal actuator was constructed. The dynamic effects are currently being examined.

Development of a Rotating Detonation Engine Facility at the University of Cincinnati

A new, air-breathing rotating detonation engine (RDE) facility has been constructed at the University of Cincinnati. This facility is built upon the modular designs currently in use at the AFRL, developed by Shank et al. in 2012 and was constructed as part of an ongoing research effort into air-breathing detonation engines. Operation of the facility has been achieved for hydrogen-air mixtures with the potential to expand to a variety of gaseous fuels. High speed and low speed instrumentation provide time-resolved wave speed and combustor pressure to characterize operation. The facility will be used to expand and improve RDE diagnostics, and establish and characterize operation for various injection schemes, fuel blends, and chamber pressures.

ACTIVE FLOW CONTROL OF SEPARATION ON A WING WITH OSCILLATORY CAMBER

Active ∞ow control of a wing with a oscillatory camber is discussed. The wing uses an adaptive actuator mounted internally to alter the shape of the suction surface which results in a change in the efiective camber by increasing the maximum thickness and moving the location of maximum thickness aft. Since the actuator motion can be altered continuously, this allows the wing shape to be either static or dynamic. Basic control options are discussed for the general case and preliminary results and observations are presented. Dynamic oscillations are used to control the separation over the wing. Experiments in a wind tunnel at low speeds are conducted using PIV and ∞ow visualization. These are compared to results from a numerical investigation. c Chord length Rec Reynolds number based on length c f Actuation frequency, Hz f + Reduced frequency, f ¢c=U FFP Forward ∞ow probability U Freestream velocity fi Angle of attack, deg

FLOW AND ACOUSTIC RADIATION FROM REALISTIC TACTICAL JET C-D NOZZLES

This paper describes a study of sound generated by supersonic jets such as those emanating from the exhausts of high-performance military aircraft. Four different nozzles are employed: one convergent nozzle and three convergent-divergent nozzles with design Mach numbers of 1.3, 1.5 and 1.65. All four nozzles have been built as laboratory-scale test articles and large-eddy simulations have been applied to jet flows from the C-D nozzle with a design Mach number of 1.5. They are all studied at their design Mach number and at an underexpanded condition. The convergentdivergent nozzles are also studied at an overexpanded condition. CFD and shadowgraph both show that these nozzles produce a shock at the nozzle throat which is present whether the nozzle is operated at design or offdesign. Acoustic measurements also show that shock-associated noise is generated even at the design condition. Examination of the curve of OASPL as a function of fully-expanded Mach number has found no minimum at or near the design Mach number.

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.

Micro-jet flow control for noise reduction of a supersonic jet from a practical C-D nozzle

A convergent-divergent nozzle similar to those used on high performance tactical aircraft has been built with micro jets at the trailing edge. These micro jets are used to increase mixing in the shear layer between the core flow and a secondary flow. Near field pressure measurements, far field acoustic measurements, as well as PIV measurements have been taken at a core flow Mach number of 1.56 and at a secondary flow Mach numbers of 0.0, 0.1, and 0.3. These results are compared with a baseline nozzle with the same geometry to see what affect the micro jets have on the jet noise production of the nozzle. A numerical simulation of the micro jets has been done using a large eddy simulation. The numerical results are compared with the experimental results. The micro jets are shown to decrease the sound pressure levels for low frequencies and increase them for high frequencies. These effects become more prominent as the secondary flow of Mach number increases. The presence of the micro jets decreases the OASPL of the core flow for most cases. The azimuth direction that benefits most changes from forward observation angles to aft observation angles as the secondary flow increases.

ACOUSTIC EFFECT of CHEVRONS on JETs Exiting Conical C-D NOZZLES

This paper describes a joint study of the acoustics of conical convergent-divergent (C-D) nozzles such as those found on high-performance military aircraft. In this paper we examine the influence of chevrons on a nozzle with an area ratio corresponding to a design Mach number of 1.5 (design pressure ratio of 3.67). The nozzle is tested at its design condition and at pressure ratios of 2.5, 3.0 and 3.5 representing overexpanded conditions and 4.0, 4.5 and 5.0 representing underexpanded conditions. Each case is compared to a baseline nozzle at the same condition without chevrons. Shadowgraph images show that chevrons reduce the shock cell spacing, and introduce additional weak shock waves. Far-field acoustic measurements show that the application of chevrons reduces screech, broad-band shock-associated noise and mixing noise except for frequencies above the broad-band shock-associated noise peak for over-expanded and perfectly expanded conditions. Near-field acoustic reveal that chevrons produce significant noise near the nozzle exit across the whole range of frequencies, but reduce mixing noise, and broad-band shock-associated noise elsewhere and they all but eliminate screech.

Fuel Blending as a Means to Achieve Initiation in a Rotating Detonation Engine

A fuel blending technique is proposed to achieve detonation initiation of hydrocarbon-air mixtures in a rotating detonation engine (RDE). An experimental investigation is performed at the University of Cincinnati Detonation Engine Test Facility to demonstrate the efficacy of gradual fuel transition to establish rotating detonation for less detonable hydrocarbons. Baseline ethylene operability is hindered by the existence of strong, low frequency instability, generally resulting in prompt failure of the detonation. An investigation of hydrogen-ethylene fuel blends does not exhibit an expanded lightoff range, but does yield significantly augmented detonation stability and improved recovery from extinction events related to the instability. Attempts to transition from a hydrogen-assisted state to an ethylene-air mixture are hindered by the immediate onset of instability in the absence of hydrogen. As the ZND analysis predicts a limited chemical effect for the range of hydrogen fractions explored in this study, the stabilizing mechanism of hydrogen addition may be physical rather than chemical, and is possibly due to suppression of fuel plenum feedback.